basic

EASA Part-66 Aviation Maintenance Maintenance Technician Technician Certification Series NO COST REVISION/UPDATE SUBSCRIPTION PROGRAM

Complete EASA Part-66 Aviation Maintenance Technician Certification Series

NO COST REVISION/UPDATE PROGRAM Aircraf Airc raf t Technica l B ook Comp Company any is offe ring a revi revision sion/update /update prog program ram to our o ur custom ers who purc purchase hase an EASA Module from the EASA Aviation Maintenance Technician Certification Series. The update is good for two (2) years from time of registration of any EASA Module or EASA bundled kits. If a revision occurs within two (2) years from date of registration, we will send you the revised pages FREE of cost to the registered email. Go to the link provided at the bottom of this page and fill out the form to be included in the EASA Revision/Update Subscription Program. In an effort to provide quality customer service please let us know if your email you register with changes so we can update our records. If you have any questions about this process please send an email to: [email protected] HERE’S HOW IT WORKS 1. All EASA Module Series textbooks textbooks contain an EASA subscription page explaining the subscription update process and provide a web site link to register for the EASA Revision/Update Subscription Program. 2. Go to the link provided below and fill out the web based form with your first and last name, current email if applicable. address, add ress, a nd school sc hool if 3. 4. 5.

From the time of of purchase, if a revision occurs to the the Module you have registered for, for, a revised PDF file containing the pages with edits will be sent to the registered email provided. Please note that that we try to keep our records as current as possible. If your email address provided at time of registration changes please let us know as soon as possible so we can update your account. This service is FREE of charge for two (2) years from date of registration.

LINK TO REGISTER FOR REVIS ION/UPD ION/UPDA ATE PROGRAM

http:///www. http: /www.actechbooks actechbooks.com/ .com/easasub/ easasub/m08 m08

MODULE 08 FOR B1 & B2 CERTIFICATION

BASIC AERODYNAMICS

Aviation Maintenance Technician Technician Certification Series

72413 U.S. Hwy 40 Tabernash, CO 80478-0270 USA www.actechbooks.com +1 970 726-5111

AVAILABLE IN

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AVIATION MAINTENANCE TECHNICIAN CERTIFICATION SERIES Author Layout/Design

Thomas Forenz Michael Amrine

Copyright © 2016 — Aircraft Technical Book Company. All Rights Reserved.

No part of this publication may be reproduced, stored in a retrieval system, transmitted in any form or by any means, electronic, mechanical, photocopying, recording or otherwise, without the prior written permission of the publisher.

To order books or for Customer Service, please call +1 970 726-5111.

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Printed in the United States of America

WELCOME Te publishers of this Aviation Maintenance echnician Certification Series welcome you to the world of aviation maintenance. As you move towards EASA certification, you are required to gain suitable knowledge and experience in your chosen area. Qualification on basic subjects for each aircraft maintenance license category or subcategory is accomplished in accordance with the following matri x. Where applicable, subjects are indicated by an "X" in the column below the license heading. For other educational tools created to prepare candidates for licensure, contact Aircraft echnical Book Company. We wish you good luck and success in your studies and in your aviation career!

EASA LICENSE CATEGORY CHART

Module number and title

A1 Airplane Turbine

B1.1 Airplane Turbine

B1.2 Airplane Piston

B1.3 Helicopter Turbine

B2 Avionics

1

Mathematics

X

X

X

X

X

2

Physics

X

X

X

X

X

3

Electrical Fundamentals

X

X

X

X

X

4

Electronic Fundamentals

X

X

X

X

5

Digital Techniques / Electronic Instrument Systems

X

X

X

X

X

6

Materials and Hardware

X

X

X

X

X

7A

Maintenance Practices

X

X

X

X

X

8

Basic Aerodynamics

X

X

X

X

X

9A

Human Factors

X

X

X

X

X

10

Aviation Legislation

X

X

X

X

X

11A

Turbine Aeroplane Aerodynamics, Structures and Systems

X

X

11B

Piston Aeroplane Aerodynamics, Structures and Systems

12

Helicopter Aerodynamics, Structures and Systems

13

Aircraft Aerodynamics, Structures and Systems

X

14

Propulsion

X

15

Gas Turbine Engine

16

Piston Engine

17A

Propeller

X X

X

X

X X

X

X

X

FORWARD PAR-66 and the Acceptable Means of Compliance (AMC) and Guidance Material (GM) of the European Aviation Safety Agency (EASA) Regulation (EC) No. 1321/2014, Appendix 1 to the Implementing Rules establishes the Basic Knowledge Requirements for those seeking an aircraft maintenance license. Te information in this Module of the Aviation Maintenance echnical Certification Series published by the Aircraft echnical Book Company meets or exceeds the breadth and depth of knowledge subject matter referenced in Appendix 1 of the Implementing Rules. However, the order of the material presented is at the discretion of the editor in an effort to convey the required knowledge in the most sequential and comprehensible manner. Knowledge levels required for Category A1, B1, B2, and B3 aircraft maintenance licenses remain unchanged from those listed in Appendix 1 Basic Knowledge Requirements. ables from Appendix 1 Basic Knowledge Requirements are reproduced at the beginning of each module in the series and again at the beginning of each Sub-Module. How numbers are written in this book: Tis book uses the International Civil Aviation Organization (ICAO) standard of writing numbers. Tis method displays large numbers by adding a space between each group of 3 digits. Tis is opposed to the American method which uses commas and the European method which uses periods. For example, the number one million is expressed as so: ICAO Standard European Standard American Standard

1 000 000 1.000.000 1,000,000

SI Units: Te International System of Units (SI) developed and maintained by the General Conference of Weights and Measures (CGPM) shall be used as the standard system of units of measurement for all aspects of international civil aviation air and ground operations. Prefixes: Te prefixes and symbols listed in the table below shall be used to form names and symbols of the decimal multiples and submultiples of International System of Units (SI) un its. MULTIPLICATION FACTOR

1 000 000 000 000 000 000 1 000 000 000 000 000 1 000 000 000 000 1 000 000 000 1 000 000 1 000 100 10 0.1 0.01 0.001 0.000 001 0.000 000 001 0.000 000 000 001 0.000 000 000 000 001 0.000 000 000 000 000 001

= 10�⁸ = 10�⁵ = 10�� = 10⁹ = 10⁶ = 10� = 10� = 10� =10�� = 10�� = 10�� = 10�⁶ = 10�⁹ = 10�� = 10��⁵ = 10��⁸

PREFIX

SYMBOL

exa peta tera giga mega kilo hecto deca deci centi milli micro nano pico femto atto

E P T G M k h da d c m µ n p f a

International System of Units (SI) Prefixes

PREFACE

Tis module includes an examination of the discipline known as aerodynamics. Even though a technician may never fly in an aircraft, the aircraft upon which he or she works conforms to the rules of aerodynamics and it is the technician's responsibility to ensure it does so at all times. Very subtle changes to the physical structure of the aircraft can cause dramatic changes in flight characteristics. Te EASA certified maintenance professional must be familiar with the basic aspects of aerodynamics to exercise the best judgment of aircraft air worthiness and aerodynamic functioning.

Module 08 Syllabus as outlined in PAR-66, Appendix 1. LEVELS CERTIFICATION CATEGORY

¦

Sub-Module 01 - Physics of the Atmosphere International Standard Atmosphere (ISA), application to aerodynamics. Sub-Module 0 2 - Aerodynamics Airflow around a body; Boundary layer, laminar and turbulent flow, free stream flow, relative airflow, upwash and downwash, vortices, stagnation; Te terms: camber, chord, mean aerodynamic chord, profile (parasite) drag, induced drag, center of pressure, angle of attack, wash in and wash out, fineness ratio, wing shape and aspect ratio; Trust, Weight, Aerodynamic Resultant; Generation of Lift and Drag: Angle of Attack, Lift coefficient, Drag coefficient, polar curve, stall; Aerofoil contamination including ice, snow, frost. Sub-Module 03 - Teory of Flight Relationship between lift, weight, thrust and drag; Glide ratio; Steady state flights, performance; Teory of the turn; Influence of load factor: stall, flight envelope and structural limitations; Lift augmentation. Sub-Module 04 - Flight Stability and Dynamics Longitudinal, lateral and directional stability (active and passive).

Module 08 - Basic Aerodynamics

B1

B2

2

2

2

2

2

2

2

2

v

REVISION LOG VERSION

ISSUE DATE

001 002

2015 01 2016 07

DESCRIPTION OF CHANGE

Module Creation and Release Minor Style Revisions and Reissue

MODIFICATION DATE

2016 07

ACKNOWLEDGMENTS

vi


CONTENTS

BASIC AERODYNAMICS Welcome ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Forward‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Preface‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Revision Log ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Acknowledgments ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥

iii iv v vi vi

SUB-MODULE 01 PHYSICS OF THE ATMOSPHERE

Knowledge Requirements ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Basic Aerodynamics ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Physics of the Atmosphere ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Pressure ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Density ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Humidity ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ emperature and Altitude ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ International Standard Atmosphere ‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Questions ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Answers ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥

1.1 1.2 1.2 1.2 1.4 1.4 1.4 1.5 1.7 1.8

SUB-MODULE 02 AERODYNAMICS

Knowledge Requirements ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Aerodynamics ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Velocity and Acceleration ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Newton’s Laws of Motion ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Bernoulli’s Principle and Subsonic Flow ‥‥‥‥‥‥‥‥‥‥‥ Boundary Layer and Friction Effects‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Planform and Vortices ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Aerodynamic erms ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Airfoils ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Shape of the Airfoil ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Airfoil Contamination‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Trust, Weight, Lift and Drag ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Aerodynamic Resultant ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Generation of Lift and Drag ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Lift and Drag Coefficients ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Lift/Drag Ratio ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Angle of Attack (AOA) ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Drag ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Parasite Drag ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Form Drag ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Friction Drag‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Interference Drag ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Induced Drag ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Wave Drag ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Drag and Airspeed‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Module 08 - Basic Aerodynamics

2.1 2.2 2.2 2.2 2.3 2.4 2.7 2.7 2.8 2.8 2.9 2.10 2.10 2.11 2.11 2.11 2.12 2.14 2.14 2.14 2.14 2.15 2.15 2.16 2.17

Questions ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ 2.19 Answers ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ 2.20

SUB-MODULE 03 THEORY OF FLIGHT

Knowledge Requirements ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Relationship Between Lift, Weight, Trust and Drag Weight, Center of Gravity, Lift ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Lift ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Trust and Drag ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Glide Ratio ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Polar Curve ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Aerodynamic Forces in urns ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Stalls ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Flight Envelope and Structural Limitations ‥‥‥‥‥‥‥‥ Load Factors ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Load Factors in Aircraft Design ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ High Speed Flight ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Subsonic vs Supersonic Flow ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Speed Ranges ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Mach Number vs Airspeed ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Shock Waves ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Sweepback ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Lift Augmentation ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Flaps ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Leading Edge Devices ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Fixed Airflow Devices ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Questions ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ Answers ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥

3.1 3.2 3.2 3.2 3.3 3.3 3.5 3.5 3.7 3.9 3.9 3.9 3.10 3.10 3.11 3.12 3.13 3.14 3.16 3.16 3.17 3.18 3.19 3.20

SUB-MODULE 04 FLIGHT STABILITY AND DYNAMICS

Knowledge Requirements ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ 4.1 Te Axes of an Aircraft ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ 4.2 Stability and Control ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ 4.3 Static Stability ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ 4.3 Longitudinal Stability (Pitching) ‥‥‥‥‥‥‥‥‥‥‥‥‥‥ 4.4 Lateral Stability (Rolling) ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ 4.7 Dihedral ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ 4.7 Sweepback ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ 4.7 Keel Effect and Weight Distribution ‥‥‥‥‥‥‥‥‥ 4.7 Vertical (Directional) ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ 4.8 Stability (Yawing) ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ 4.8 Free Directional Oscillations (Dutch Roll) ‥‥‥‥‥‥ 4.9 Questions ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ 4.11 Answers ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ 4.12

vii

CONTENTS

Acronym Index ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ A.1 Glossary ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ G.1 Index ‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥‥ I.1

viii


E E H R T E F H O P S S C I O S M Y T A H P

PART-66 SYLLABUS LEVELS CERTIFICATION CATEGORY

¦

B1

B2

2

2

Sub-Module 01

PHYSICS OF THE ATMOSPHERE

Knowledge Requirements 8.1 - Physics of the Atmosphere International Standard Atmosphere (ISA), application to aerodynamics.

Level 2

A A general kknowledge nowledge of the the ttheoretical heoretica l and pract practical ical aspects of tthe he subject and an ability to apply that knowledge. Objectives: (a) Te applicant should be able to understa nd the theoretic al fundamentals of the subject. (b) Te applicant should be able to give a general descript description ion of the subject using, as appropriate, typical examples. (c) Te applicant should be able to use mathematic al formula in mathematical conjunction with physical laws describing the subject. (d) Te applicant should be able to read and understa nd sketche sketches, s, drawings and schematics describing the subject. (e) Te applicant should be able to apply his knowledge in a practic al manner using detailed procedures.


1.1

BASIC AERODYNAMICS Tree topics that are directly related to the manufacture, operation, and repair of aircraft are: aerodynamics, aircraft assembly, and rigging. Each of these subject areas, though studied separately, eventually connect to provide a scientific and physical understanding of how an aircraft is prepared for flight. A logical place to start with these three topics is the study of basic aerodynamics. By studying aerodynamics, a person becomes familiar with the fundamentals of aircraft flight. Aerodynamics is the study of the dynamics of gases. Te

interaction between a moving object and the atmosphere is the primary interest in this module. Te movement of an object and its reaction to the air flow around it can be seen when watching water passing the bow of a ship. Te major difference between water and air is that air is compressible and water is incompressible. Te action of the airflow over a body is a large part of the study of aerodynamics. Some common aircraft terms, such as rudder, hull, water line, and keel beam, were borrowed from nautical terms. Many textbooks have been written about the aerodynamics of aircraft flight. It is not necessary for an airframe and powerplant technician to be as knowledgeable as an aeronautical design engineer

about aerodynamics. he technician must be able to understand the relationships between how an aircraft performs in flight and its reaction to the forces acting on its structural parts. Understanding why aircraft are designed with particular types of primary and secondary control systems and why the surfaces must be aerodynamically smooth becomes essential when maintaining today’s complex aircraft. Te theory of flight should be described in terms of the laws of flight because what happens to an aircraft when it flies is not based upon assumptions, but upon a series of facts. Aerodynamics is a study of laws which have been proven to be the physical reasons why an airplane flies. he term aerodynamics is derived from the combination of two Greek words: "aero," meaning air, and "dyne," meaning force of power. Tus, when "aero" joins "dynamics" the result is "aerodynamics "—the study of objects in motion through the air and the forces that produce or change such motion. Aerodynamically, an aircraft can be defined as an object traveling through space that is affected by the changes in atmospheric conditions. o state it another way, aerodynamics covers the relationships bet ween the aircraft, relative wind, and atmosphere.

PHYSICS OF THE ATMOSPHERE Before examining the fundamental laws of flight, several basic facts must be considered. An aircra ft operates in the air. Terefore, those properties of air that affect the control and performance of an aircraft must be understood. Te air in the earth’s atmosphere is composed mostly of nitrogen and oxygen. Air is considered a fluid because it fits the definition of a substance that has the ability to flow or assume the shape of the container in which it is enclosed. If the container is heated, pressure increases; if cooled, the pressure decreases. he weight of air is heav iest at sea le vel where it has been compressed by all of the air above. his compression of air is called atmospheric pressure.

1.2

PRESSURE Atmospheric press ure is usually defined as the force

exerted against the earth ’s surface by the weight of the air above that surface. Weight is force applied to an area that results in pressure. Force (F) equals area (A) times pressure (P), or F = AP. Terefore, to find the amount of pressure, divide area into force (P = F/A). A column of air (one square inch) extending from sea level to the top of the atmosphere weighs approximately 14.7 pounds; therefore, atmospheric pressure is stated in pounds per square inch (psi). Tus, atmospheric pressure at sea level is 14.7 psi. ( Figure 1-1) Atmospheric pressure is measured with an instrument called a barometer, composed of mercury in a tube that records atmospheric pressure in inches of mercury (Hg). ( Figure 1-2)


Thermosphere

Height of Earth’s atmosphere (about 50 miles) h c n i 1


350 km

1 i n c h

Mesosphere Entire air column weight 14.7lb

90 km

Ozone layer

50 km

Stratosphere

Tropopause

18 km

Troposphere

14 km

Sea level Earth

Figure 1-1. The weight exerted by a 1 square inch column of air stretching from sea level to the top of the atmosphere is what is measured when it is said that atmospheric pressure is equal to 14.7 pounds per square inch.

Standard Sea Level Pressure

29.92" Hg

Inches of Mercury 30

m u u c a V

Millibars

Standard Sea Level Pressure

1016

25

847

20

677

15

508

10

339

5

170

0

0

Te standard measurement in aviation altimeters and U.S. weather reports has been "Hg". However, world wide weather maps and some non-U.S., manufactured aircraft instruments indicate pressure in millibars (mb), an SI metric unit.

1013

mb

Aviators often interchange references to atmospheric pressure between linear d isplacement (e.g., inches of mercury) and units of force (e.g., psi). Over the years, meteorology has shifted its use of linear displacement representation of atmospheric pressure to units of force. Te unit of force nearly universally used today to represent atmospheric pressure in meteorology is the hectopascal (hPa). A pascal is a SI metric unit that expresses force in Newtons per square meter. A hectoPascal is 100 Pascals. 1 013.2 hPa is equal to 14.7 psi which is equal to 29.92 Hg. ( Figure 1-3)

1"

1"

" 1

0.491 lb Mercury

Figure 1-2. Barometer used to measure atmospheric pressure.

Atmospheric Pressure Standard atmospheric pressure at sea level is also known as 1 atmosphere, or 1 atm. The following measurements of standard atmospheric pressure are all equal to each other. 1 atm (atmosphere)

=

14.7 psi (prounds per square inch)

=

29.92 in Hg (inches of mercury)

=

1013.2 hPa (or 101325 newtons per square meters)

=

1013.2 mb (millibars)

=

760 mm Hg (millimeters of mercury)

Figure 1-3. Various equivalent representations of atmospheric pressure at sea level. Module 08 - Basic Aerodynamics

1.3

Atmospheric pressure decreases with increasing altitude. Te simplest explanation for this is that the column of air that is weighed is shorter. How the pressure changes for a given altitude is shown in Figure 1-4 . Te decrease in pressure is a rapid one and, at 50 000 feet, the atmospheric pressure has dropped to almost one-tenth of the sea level value. As an aircraft ascends, atmospheric pressure drops, the quantity of oxygen decreases, and temperature drops. Tese changes in altitude affect an aircraft’s performance in such areas as lift and engine horsepower. Te effects of temperature, altitude, and density of air on aircraft performance are covered in the following paragraphs.

100 000 80 000 ) t e e f ( e d u t i t l A

60 000 40 000 20 000 Sea level 0

2

4

6

8

10

12

14

Pressure (pounds per square inch)

Figure 1-4. Atmospheric pressure decreasing with altitude. At sea level the pressure is 14.7 psi, while at 40 000 feet, as the dotted lines show, the pressure is only 2.72 psi.

DENSITY Density is weight per unit of volume. Since air is a

mixture of gases, it can be compressed. If the air in one container is under half as much pressure as an equal amount of air in an identical container, the air under greater pressure is twice as dense as that in the other container. For the equal weight of air, that which is under the greater pressure occupies only half the volume of that under half the pressure.

Tus, air at high altitudes is less dense than air at low altitudes, and a mass of hot air is less dense than a mass of cool air. Changes in density affect the aerodynamic performance of aircraft with the same horsepower. An aircraft can fly faster at a high altitude where the air density is low than at a low altitude where the density is greater. his is because air offers less resistance to the aircraft when it contains a smaller number of air particles per unit of volume.

Te density of gases is governed by the following rules: 1. Density varies in direct proportion with the pressure. 2. Density varies inversely with the temperature.

HUMIDITY Humidity is the amount of water vapor in the air. Te

maximum amount of water vapor that air can hold varies with the temperature. Te higher the temperature of the air, the more water vapor it can absorb. 1. Absolute humidity is the weight of water vapor in a unit volume of air. 2. Relative humidity is the ratio, in percent, of the moisture actually in the air to the moisture it would hold if it were saturated at the same temperature and pressure. Assuming that the temperature and pressure remain the same, the density of the air varies inversely with the humidity. On damp days, the air density is less than on dry days. For this reason, an aircraft requires a longer runway for takeoff on damp days than it does on dry days. 1.4

By itself, water vapor weighs approximately five-eighths as much as an equal amount of perfectly dry air. herefore, when air contains water vapor, it is not as heavy as dry air containing no moisture.


TEMPERATURE AND ALTITUDE emperature variations in the atmosphere are of concern to aviators. Weather systems produce changes in temperature near the earth’s surface. emperature also changes as altitude is increased. Te troposphere is the lowest layer of the atmosphere. On average, it ranges from the earth’s surface to about 38 000 feet above it. Over the poles, the troposphere extends to only 25 000 30 000 feet and, at the equator, it may extend to around 60 000 feet. Tis oblong nature of the troposphere is illustrated in Figure 1-5 . Most civilian aviation takes place in the troposphere in which temperature decreases as altitude increases. Te rate of change is somewhat constant at about –2 °C or –3.5 °F for every 1 000 feet of increase in altitude. Te upper boundary of the troposphere is the tropopause. It is characterized as a zone of relatively constant temperature of –57 °C or –69 °F.


100 000–116 000 feet 137 000–153 000 feet 25 000–30 000 feet N Equator S

55 000–65 000 feet

T r o po sp h e r e

S tr a t osp h e r e

M es osphe r e T he rm osp he r e

Figure 1-5. The troposphere extends higher above the earth’s surface at the equator than it does at the poles.

Above the tropopause lies the stratosphere. emperature increases with altitude in the stratosphere to near 0 °C before decreasing again in the mesosphere, which lies above it. he stratosphere contains the ozone layer that protects the earth’s inhabitants from harmful UV (Ultraviolet) rays. Some civilian flights and numerous military flights occur in the stratosphere.

90

80 Thermosphere

70

60

Figure 1-6 diagrams the temperature variations in

different layers of the atmosphere. As stated, density varies inversely with temperature or, as temperature increases, air density decreases. his phenomenon explains why on very warm days, aircraft takeoff performance decreases. he air available for combustion is less dense. Air with low density contains less total oxygen to combine with the fuel.

160 000 ft

Mesopause

T em per at u r e

50

Mesosphere 40

160 000 ft

) m k ( t h g i e H

30

Stratopause

Stratosphere 20 Ozone layer 10 38 000 ft

Tropopause

Troposphere

Mt. Everest –100 –140

–80

–60

–100

–40

–60

–20 –20 0 20

0

20 60

40 50 °C 100 120 °F

Temperature

Figure 1-6. The atmospheric layers with temperature changes depicted by the red line.


1.5

INTERNATIONAL STANDARD ATMOSPHERE Te atmosphere is never at rest. Pressure, temperature, humidity, and density of the air are continuously changing. o provide a basis for theoretical calculations, performance comparisons and instrumentation parity, standard values for these and other characteristic of the atmosphere have been developed. International

ALTITUDE

Civil Aviation Organization (ICAO), International Organization for Standardization ( ISO), and various governments establish and publish the values known as the International Standard Atmosphere. ( Figure 1-7 )

TEMPERATURE

PRESSURE

DENSITY

psi

hPa

slug/ft³

kg/m³

15

14.67

1013.53

0.002378

1.23

55.4

13

14.17

977.16

0.002309

1.19

2000

51.9

11

13.66

941.82

0.002242

1.15

3000

48.3

9.1

13.17

908.11

0.002176

1.12

4000

44.7

7.1

12.69

874.94

0.002112

1.09

5000

41.2

5.1

12.05

843.07

0.002049

1.06

6000

37.6

3.1

11.78

812.2

0.001988

1.02

7000

34

1.1

11.34

781.85

0.001928

0.99

8000

30.5

-0.9

10.92

752.91

0.001869

0.96

9000

26.9

-2.8

10.5

724.28

0.001812

0.93

10 000

23.3

-4.8

10.11

697.06

0.001756

0.9

15 000

5.5

-14.7

8.3

571.82

0.001496

0.77

20 000

-12.3

-24.6

6.75

465.4

0.001267

0.65

25 000

-30.2

-34.5

5.46

376.01

0.001066

0.55

30 000

-48

-44.4

4.37

301.3

0.000891

0.46

35 000

-65.8

-54.3

3.47

238.42

0.000738

0.38

40 000

-69.7

-56.5

2.72

187.54

0.000587

0.3

45 000

-69.7

-56.5

2.15

147.48

0.000462

0.24

50 000

-69.7

-56.5

1.68

115.83

0.000362

0.19

Feet

°F

Sea Level

59

1000

°C

Figure 1-7. The International Standard At mosphere.

1.6


QUESTIONS

Question: 1-1 Atmospheric pressure is measured with an instrument called a __________________.

Question: 1-3 In which layer of the atmosphere does most civilian aviation take place?

Question: 1-2 An aircraft can fly __________________ at a high altitude where the air density is low than at a low altitude where the air density is greater.


1.7

ANSWERS

Answer: 1-1 barometer.

Answer: 1-3 roposphere.

Answer: 1-2 faster.

1.8



¦

B1

B2

2

2

Sub-Module 02

AERODYNAMICS

Knowledge Requirements 8.2 - Aerodynamics Airflow around a body; Boundary layer, laminar and turbulent flow, free stream flow, relative airflow, upwash and downwash, vortices, stagnation; Te terms: camber, chord, mean aerodynamic chord, profile (para site) drag, induced drag, center of pressure, angle of attack, wash in and wash out, fineness ratio, wing shape and aspect ratio; Trust, Weight, Aerodynamic Resu ltant; Generation of Lift and Drag: Angle of Attack, Lift coefficient, Drag coefficient, polar curve, stall; Aerofoil contamination including ice, snow, frost.

Level 2

A general knowledge of the theoretical and practical aspects of the subject and an ability to apply that knowledge. Objectives:

(a) Te applicant should be able to understa nd the theoretic al fundamentals of the subject. (b) Te applicant should be able to give a general description of the subject using, as appropriate, typical examples. (c) Te applicant should be able to use mathematical formula in conjunction with physical laws describing the subject. (d) Te applicant should be able to read and understa nd sketches, drawings and schematics describing the subject. (e) Te applicant should be able to apply his knowledge in a practic al manner using detailed procedures.


2.1

S C I M A N Y D O R E A

AERODYNAMICS Te law of conservation of energy states that energy may neither be created nor destroyed. Motion is the act or process of changing place or position. An object may be in motion with respect to one object and motionless with respect to another. For example, a person sitting quietly in an aircraft flying at 200 knots is at rest or motionless with respect to the aircraft; however, the person and the aircraft are in motion with respect to the air and to the earth.

Air has no force or power, except pressure, unless it is in motion. When it is moving, however, its force becomes apparent. A moving object in motionless air has a force exerted on it as a result of its own motion. It makes no difference in the effect then, whether an object is moving with respect to the air or the air is moving with respect to the object. Te flow of air around an object caused by the movement of either the air or the object, or both, is called the relative wind.

VELOCITY AND ACCELERATION Te terms speed and velocity are often used interchangeably but they do not have the same meaning. Speed is the rate of motion in relation to time, and velocity is the rate of motion in a particular direction in relation to time. An aircraft starts from New York City and flies 10 hours at an average speed of 260 kilometers per hour (kph). At the end of this time, the aircraft may be over the Atlantic Ocean, Canada the Gulf of Mexico, or, if its flight were in a circular path, it may even be back over New York City. If this same aircraft flew at a velocity of

260 kph in a southwestward direction, it would arrive in Dallas, X in about 10 hours. Only the rate of motion is indicated in the first example and denotes the speed of the aircraft. In the last example, the particular direction is included with the rate of motion, thus, denoting the velocity of the aircraft. Acceleration is defined as the rate of change of velocity. An

aircraft increasing in velocity is an example of positive acceleration, while another aircraft reducing its velocity is an example of negative acceleration, or deceleration.

NEWTON’S LAWS OF MOTION Te fundamental laws governing the action of air about a wing are known as Newton’s laws of motion. Newton’s first law is normally referred to as the law of inertia. It simply states that a body at rest does not move unless force is applied to it. If a body is moving at uniform speed in a straight line, force must be applied to increase or decrease the speed. According to Newton’s law, since air has mass, it is a body. When an aircraft is on the ground with its engines off, inertia keeps the aircraf t at rest. An aircraft is moved from its state of rest by the th rust force created by a propeller, or by the expanding exhaust, or both. When an aircraft is flying at uniform speed in a straight line, inertia tends to keep the ai rcraft moving. Some external force is required to change the aircraft from its path of flight. Newton’s second law states that if a body moving wit h uniform sp ee d is ac ted up on by an externa l 2.2

force, the change of motion is proportional to the amount of the force, and motion takes place in the direction in which the force acts. his law may be stated mathematically as follows: Force = mass × acceleration (F = ma) If an aircraft is flying against a headwind, it is slowed down. If the wind is coming from either side of the aircraft’s heading, the aircraft is pushed off course unless the pilot takes corrective action against the wind direction. Newton’s third law is the law of action and reaction. Tis law states that for every action (force) there is an equal and opposite reaction (force). his law can be illustrated by the example of firing a gun. Te action is the forward movement of the bullet while the reaction is the backward recoil of the gun. Te three laws of motion that have been discussed apply to the theory of flight. In many cases, all three laws may be operating on an aircraft at the same time. Module 08 - Basic Aerodynamics

BERNOULLI’S PRINCIPLE AND SUBSONIC FLOW Bernoulli’s principle states that when a fluid (air) flowing through a tube reaches a constriction, or narrowing, of the tube, the speed of the fluid flowing through that constriction increases and its pressure decreases. he cambered (curved) surface of an airfoil (wing) affects the airflow exactly as a constriction in a tube affects airflow. ( Figure 2-2) Diagram A of Figure 2-1 illustrates the effect of air passing through a constriction in a tube. In B, air is flowing past a cambered surface, such as an airfoil, and the effect is similar to that of air passing through a restriction. An airfoil is a surface designed to obtain lift from the air through which it moves. As the air flows over the curved upper surface of an airfoil, its velocity increases and its pressure decreases; an area of low pressure is formed. Tere is an area of greater pressure on the lower surface of the airfoil, and this greater pressure tends to move the wing upward. Te difference in pressure between the upper and lower surfaces of the wing is called lift . Tree-fourths of the total lift of an airfoil is the result of the decrease in pressure over the upper surface. Te impact of air on the lower surface of an airfoil produces the other one-fourth of the total lift. Note that in order to fit the model of Bernoulli's Principle, the airflow over the wing surfaces must be laminar. Laminar air flow refers to airflow that is flowing A

in a consistent smooth stream. urbulent flow is also possible. Tis is where the air flowing over the surface no longer so closely adheres to it. Te flow is thicker and faster, however, some lift is produced. When the airflow actually separates from the surface of a wing, a different type of turbulence occurs. Tis type of turbulence does not produce lift and Bernoulli's Principle does not apply. More discussion of these phenomena occur below in the section entitled Boundary Layer and Friction Effects. Free stream airflow is air flowing without obstruction before it engages the aircraft structure. Te velocity of the free stream flow is equal to the speed aircraft. Te pressure of the free stream airflow is static pressure. When the free st ream f low arr ives at the aircra ft structure, such as the wing, it must flow around the surface areas. As it does so, the pressure and velocity of the air change depending on the shape of the wing. here is a point in front of the struct ure, however, where the velocity of the air is zero. Tis is known as the point of stagnation. ypical airflow patterns show the relationship between static pressure and velocity defined by Bernoulli. In aerodynamics, when positive pressure is mentioned, it refers to pressures above atmospheric pressure. Negative pressure or suction pressure is lower than atmospheric pressure. Any object placed in an airstream will have

Mass of Air S a m e M

Normal Pressure

i r ass o f A

Velocity Increased Pressure Decreased (Compared to Original)

Normal Pressure

B

Normal Flow

Increased Flow

Normal Flow

Figure 2-1. Bernoulli’s Principle. Module 08 - Basic Aerodynamics

2.3


the air impact or stagnate at some point near the leading edge. Te pressure at this point of stagnation will be an absolute static pressure equal to the total pressure of the airstream. In other words, the static pressure at the stagnation point will be greater than the atmospheric pressure by the amount of the dynamic pressure of the airstream. As the f low divides and proceeds around the object, the increases in local velocity produce decreases in static pressure. his procedure of flow is best illustrated by the flow patterns and pressure distributions of Figure 2-2 . Note that the "streamlines" in the diagram show the velocity of the airflow. When they are close together, high velocity exists at that point and when they are far apart, low velocity exists at that point. Te vector arrows in the diagram show the magnitude and direction of the low pressure caused by the increased velocity of the airflow.

the airfoil now provides a net force perpendicular to the airstream in the upward direct. Tis is lift. ( Figure 2-3 ) Te creation of lift is discussed in greater detail below. Point of Stagnation

Peak Suction Pressure Forward Stagnation Point

+

+

Aft Stagnation Point

-

Figure 2-2. Velocity distribution of airflow over a symmetrical airfoil (top);

Because the object in Figure 2-2 is a symmetrical airfoil, the relative airflow striking it flows above and below the airfoil in the same manner. Te pressures are the same and no lift is produced. By reshaping the airfoil or by tilting it in relation to the relative airflow, uneven flow over the upper and lower surfaces occurs. Tis causes uneven pressure above and below the airfoil which results in the creation of lift. Simply by tilting the same symmetrical airfoil, an increase in upper surface suction occurs and the decreased in velocity on the lower surface causes a decrease in lower surface suction. Also, upwash is generated ahead of the airfoil, the forward stagnation point moves under the leading edge, and a downwash is evident aft of the airfoil. (Upwash and downwash are the deflection directions of the air as it negotiates its path around the airfoil.) Te pressure distribution on

and the resulting pressure (bottom). Positive Lift

Upwash

Increased Local Velocity Downwash

Decreased Local Velocity

Figure 2-3. Uneven airflow, uneven pressure, up-wash and down-wash are all caused by tilting the airfoil in relation to the free stream airflow.

BOUNDARY LAYER AND FRICTION EFFECTS In the study of physics and fluid mechanics, a boundary layer is that layer of fluid in the immediate vicinity of a bounding surface. In relation to an aircraft, the boundary layer is the part of the airflow closest to the surface of the aircraft. In designing high-performance aircraft, considerable attention is paid to controlling the behavior of the boundary layer to minimize pressure drag and skin friction drag.

viscous nature of airflow reduces the local velocities on a surface and accounts for the drag of skin friction. Te retardation of air particles due to viscosity is greatest immediately adjacent to the surface. At the very surface of an object, the air particles are slowed to a relative velocity of near zero. Above this area other particles experience successively smaller retardation until finally, at some distance above surface, the local velocity reaches the full value of the airstream above the surface.

Because air has viscosity (internal resistance to flow), air encounters resistance to flow over a surface. he 2.4


his layer of air over the surface which shows local retardation of airflow from viscosity is the boundary layer. he characteristics of this boundary layer are illustrated in Fig ure 2-4 with the flow of air over a smooth flat plate. Te beginning flow on a smooth surface gives evidence of a very thin boundary layer with the flow occurring in smooth laminations. Te boundary layer flow near the leading edge is similar to layers or laminations of air sliding smoothly over one another. he term for this type of flow is the "laminar" boundary layer as mentioned previously. Tis smooth laminar flow exists without the air particles moving from a given elevation above the surface. As the flow continues back from the leading edge, friction forces in the boundary layer continue to dissipate energy of the airstream and the laminar boundary layer increases in thickness with distance from the leading edge. After some distance back from the leading edge, the laminar boundary layer begins an oscillatory disturbance which is unstable. A waviness occurs in the laminar boundary layer which u ltimately grows larger and more severe and destroys the smooth

laminar flow. Tus, a transition takes place in which the laminar boundary layer decays into a "turbulent" boundary layer. he same sort of transition can be noticed in the smoke from a cigarette in still air. At, first, the smoke ribbon is smooth and laminar, then it develops a definite waviness and decays into a random turbulent smoke pattern. As soon as the transition to the turbulent boundary layer takes place, the boundary layer thickens and grows at a more rapid rate. (Te small scale, turbulent flow within the boundary layer should not be confused with the large scale turbulence associated with airflow separation.) Te flow in the turbulent boundary layer allows the air particles to travel from one layer to another producing an energy exchange. However, some small laminar flow continues to exist in the very lower levels of the turbulent boundary layer and is referred to as the "laminar sub-layer." Te turbulence which exists in the turbulent boundary layer allows determination of the point of transition by several means. Since the t urbulent boundary layer transfers heat more easily than the laminar layer, frost, water, and oil films will be removed more rapidly from Turbulant Boundary Layer

Transition region Laminar Boundary Layer

Laminar Sub Layer

Comparison of Velocity Profiles for Laminar and Turbulant Boundary Layers Turbulant Profile Laminar Profile

- Low Thickness - Low Velocities Next to Surface - Gradual Velocity Change - Low Skin Friction

- GreaterThickness - Higher Velocities Next to Surface - SharpVelocity Change - Higher Skin Friction

Figure 2-4. Boundary Layer Characteristics. Module 08 - Basic Aerodynamics

2.5


the area aft of the transition point . Also, a small probe may be attached to a stethoscope and positioned at various points along a surface. When the probe is in the laminar area, a low "hiss" will be heard. When the probe is in the turbulent area, a sharp "crackling" will be audible. In order to compare the characteristics of the laminar and turbulent boundary layers, the velocity profiles (the variation of boundary layer velocity with height above the surface) should be compared under conditions which could produce either laminar or turbulent flow. Te typical laminar and turbulent profiles are shown in Figure 2-4 . he velocity profile of the turbulent boundary layer shows a much sharper initial change of velocity but a greater height (or boundary layer thickness) required to reach the free stream velocity.

As a result of these differences, a comparison shows: 1. Te turbulent boundary layer has a fuller velocity profile and has higher local velocities immediately adjacent to the surface. Te turbulent boundary layer has higher kinetic energy in the airflow next to the surface. 2. At the surface, the laminar boundary layer has the less rapid change of velocity with distance above the surface. Since the shearing stress is proportional to the velocity gradient, the lower velocity gradient of the laminar boundary layer is evidence of a lower friction drag on the surface. In conditions of flow where a turbulent and a laminar boundary layer can exist, the laminar skin friction is about one-third that for turbulent flow. And while the low friction drag of the laminar boundary layer is desirable, the transition to turbulent boundary layer flow is natural and largely inevitable.

Tapered Leading Edge, Straight Trailing Edge

Tapered Leading and Trailing Edges

Delta Wing

Sweptback Wings

Rectangular Wing

Straight Leading Edge, Tapered Trailing Edge

Figure 2-5. Various wing planforms.

2.6


PLANFORM AND VORTICES Te previous discussion of aerodynamic forces concerned the properties of airfoil sections in two-dimensional flow with no consideration given to the influence of the planform . Te planform is the shape or outline of an aircraft wing as projected onto a horizontal plane. ( Figure 2-5 ) When the effects of wing planform are introduced, attention must be directed to the existence of flow components in the span-wise direction. In other words, the airfoil section properties considered thus far deal with flow in two dimensions. Planform properties consider flow in three dimensions. Te pressure above the wing is less than atmospheric pressure, and the pressure below the wing is equal to or greater than atmospheric pressure. Since fluids always move from high pressure toward low pressure, in addition to the movement of air over the wing from front to rear, there is also a spanwise movement of air from the bottom of the wing outward from the fuselage and upward around the wing tip. Tis flow of air results in spillage over the wing tip, thereby setting up a whirlpool of air called a " vortex ." ( Figure 2-6 ) Te plural of vortex is vortices. As the difference in the pressure between the air on the bottom and top of the wing increases, more lift is generated. his increased pressure differential also causes more violent vortices. Small aircraft pilots must be especially vigilant when flying behind large aircraft. Te vortices coming off the wingtips of a transport category aircraft could cause loss of control if encountered before they have had time to dissipate into the atmosphere.

Note that the air on the upper surface of the wing planform has a tendency to move in toward the fuselage and off the trailing edge as shown by the blue arrows in Figure 2-6. Tis air current forms a similar vortex to a wingtip vortex but at the inner portion of the trailing edge of the wing. All vortices increase drag because of the turbulence produced, and constitute induced drag. Vortices increase as lift (and drag) increase. Drag will be discussed in further detail later in this module. Just as lift increases by increasing of the angle of the airfoil into the wind, drag also increases as the angle becomes greater. Tis occurs because, within limits, as the angle is increased, the pressure difference between the top and bottom of the wing becomes greater. Tis causes more violent vortices to be set up, resulting in more turbulence and more induced drag.

t ex V or

Figure 2-6. Wingtip vortices.

AERODYNAMIC TERMS Before continuing the discussion on aerodynamics, some terms are defined and illustrations considered. Te chord of a wing is the width of the wing from the leading edge apex to the trailing edge. A chord line is a line depicting the chord which extends forward of the leading edge. It is used for angular reference to the chord. ( Figure 2-9 ) Te average chord is the area of the Module 08 - Basic Aerodynamics

wing divided by the wing span. Te mean aerodynamic chord is the average distance from the leading edge to the trailing edge of the wing. Due to the many wing planform designs, the mean aerodynamic chord is not necessarily half way from the fuselage to the wing tip as it is on a perfectly rectangular wing. However, the mean aerodynamic chord has half of the surface area 2.7


Angle of Incidence

Longitudinal Axis

C ho r d Li ne o f W in g

Figure 2-8. Angle of incidence.

in order to provide better aerodynamic characteristics especially stall characteristics. When a wing is twisted down at the tip so that the angle of incidence is less at the wingtip than it is at the wing root, it is called washout. If the wing is twisted in the opposite direction so that the wing tip angle of incidence is greater than at the wing root, it's called wash in.

MAC

Figure 2-7. Mean aerodynamic chord (MAC).

of the wing on each side of it. ( Figure 2-7 ) Te mean aerodynamic chord is used by aerodynamicists when calculating stability and other design factors. he ac ut e ang le the win g chord ma kes wit h the longitudinal axis of the aircraft is called the angle of incidence, or the angle of wing setting. ( Figure 2-8 )

Refer to Figure 2-9 to clarify the following terms. Te camber of a wing is the curve of the upper wing surface. Te lower surface of the wing also has camber. Te mean camber line lies within the wing half way between the upper camber and the lower camber. Maximum camber is located where the mean camber line is the greatest distance from the chord line.

he angle of incidence in most cases is a fixed, built-in angle. When the leading edge of the wing is higher than the trailing edge, the angle of inc idence is said Location of to be positive. he angle of Max. Thickness incidence is negative when Leading Edge the leading edge is lower than Radius the trailing edge of the wing. Chord Line Other unique features of wings inc lude wash in and wash out. A wing does not have to be constructed flat in a single plain. A wing may be twisted from root to tip

Maximum Thickness Upper Surface Maximum Camber

Mean Camber Line

Chord Line

Chord Line

Lower Surface

Leading Edge

Trailing Edge

Chord Location of Maximum Camber

Figure 2-9. Chord and camber of a wing.

AIRFOILS Since an airfoil is a surface designed to obtain lift from the air through which it moves, it can be stated that any part of the aircraft that converts air resistance into lift is an airfoil. he profile of a conventional wing is an excellent example of an airfoil. ( Figure 2-10 ) Notice that the top surface of the wing profile has greater curvature than the lower surface.

2.8

he difference in curvat ure of the upper and lower surfaces of the wing creates the lifting force. Air flowing over the top surface of the wing must reach the trailing edge of the wing in the same amount of time as the air flowing under the wing. o do this, the air passing over the top surface moves at a greater velocity than the air passing below the wing because of the greater distance it must travel along the top surface. his increased velocity, according to Bernoulli’s Principle, means a corresponding decrease in pressure on the upper surface. Module 08 - Basic Aerodynamics

115 mph

14.54 lb/in² Early Airfoil

Later Airfoil

100 mph

14.7 lb/in²

105 mph

14.67 lb/in²

Clark 'Y' Airfoil (Subsonic)


Figure 2-10. Airflow over a wing section.

Tus, a pressure differential is created between the upper and lower surfaces of the wing, forcing the wing upward in the direction of the lower pressure. SHAPE OF THE AIRFOIL

Individual airfoil section properties differ from those properties of the entire wing or aircraft as a whole because of the effect of the wing planform. A wing may have various airfoil sections from root to tip, with taper, twist, and sweepback. Te resulting aerodynamic properties of the wing are determined by the action of each section along the span. he shape of the ai rfoil determines the amount of turbulence or skin friction that it produces, consequently affecting the efficiency of the wing. urbulence and skin friction are controlled mainly by the fineness ratio, which is defined as the ratio of the chord of the airfoil to its maximum thickness. If the wing has a high fineness ratio, it is a very thin wing. A thick wing has a low fineness ratio. A wing with a high fineness ratio produces a large amount of skin friction. A wing with a low fineness ratio produces a large amount of turbulence. Te best wing is a compromise between these two extremes to hold both turbulence and skin friction to a minimum. Figure 2-11 illustrates a wide variety of airfoil shapes.

High-lift wings and high-lift devices for wings have been developed by shaping the airfoils to produce the desired effect. Te amount of lift produced by an airfoil increases with an increase in wing camber. As stated, camber refers to the curvature of an airfoil surface above and below the chord line. Upper camber refers to the upper surface, lower camber to the lower surface, and mean camber to the mean line of the section. Camber is positive when departure from the chord line is outward and negative when it is inward. Tus, high-lift wings have a large positive camber on the upper surface and a slightly negative camber on the lower surface. Module 08 - Basic Aerodynamics

Laminar Flow Airfoil (Subsonic)

Circular Arc Airfoil (Supersonic)

Double Wedge Airfoil (Supersonic)

Figure 2-11. Airfoil designs.

Wing flaps cause an ordinary wing to approximate this same condition by increasing the upper camber and by creating a negative lower camber. It is also known that the larger the wingspan, as compared to the chord, the greater the lift obtained. Tis comparison is called aspect ratio. he higher the aspect ratio, the greater the lift. In spite of the benefits from an increase in aspect ratio, there are definite limitations defined by structural and drag considerations. On the other hand, an airfoil that is perfectly streamlined and offers little wind resistance sometimes does not have enough lifting power to take the aircraft off the ground. Tus, modern aircraft have airfoils which strike a medium between extremes, the shape depending on the purposes of the aircraft for which it is designed. AIRFOIL CONTAMINATION

All discussion of aerodynamic behavior of airfoils assumes that the aircraft airfoils are free of contamination. Some of the most common forms of contamination are ice, snow and frost. Each of these, if accumulated on the aircraft, will reduce its capacity to develop lift. Ice commonly changes the shape of the airfoil which disrupts airflow and make it less efficient. Snow, ice, and especially frost, alter the smooth even surface that normally promotes laminar airflow. Laminar airflow is required to set up the pressure differential between the lower and upper wing surfaces that creates lift. All snow and ice must be completely removed from any aircraft before flight. Frost must also be removed. W hile it appears insignificant, the disruption to airflow caused by frost is possibly the most dangerous. 2.9

If ice is allowed to accumulate on the aircraft during flight ( Figure 2-12), the weight of the aircraft is increased while the ability to generate lift is decreased. As little as

0.8 millimeter of ice on the upper wing surface increases drag and reduces aircraft lift by 25 percent.

THRUST, WEIGHT, LIFT AND DRAG Tere are four forces that act upon an aircraft in flight: thrust, weight, lift and drag. A brief description of each is given. A discussion of angle of attack, lift and drag follows. Further exploration of the relationship between the four forces of flight occurs in Sub-Module 03. 1. Trust —the force that moves the aircraft forward. Trust is the forward force produced by the powerplant that overcomes the force of drag. 2. Gravity or weight —the force that pulls the aircraft toward the earth. Weight is the force of gravity acting downward upon everything that goes into the aircraft, such as the aircraft itself, crew, fuel, and cargo.

3. Lift —the force that pushes the aircraft upward. Lift acts vertically and counteracts the effects of weight. 4. Drag —the force that exerts a braking action to hold the aircraft back. Drag is a backward deterrent force and is caused by the disruption of the airflow by the wings, fuselage, and protruding objects. Figure 2-13 illustrates the vectors of these four forces.

Tey are in perfect balance only when the aircraft is in straight-and-level unaccelerated flight.

AERODYNAMIC RESULTANT An aircraft in flight is continuously affected by thrust, weight, li ft and drag. he di rect ions in which the forces act is known. Te magnitude of the forces can be calculated. When the forces are not in balance, a resultant or resulting force will exist. his is the combined force of all of the forces acting on the aircraft. In all types of flying, flight calculations are based on the magnitude and direction of the four forces. Te forces of lift and drag are the direct result of the relationship between the relative wind and the aircraft. Te force of lift always acts perpendicular to the relative wind, and the force of drag always acts parallel to and in the same direction as the relative wind. Tese forces

are actually the components that produce a resultant lift force on the wing. ( Figure 2-14 )

L i f t

D r ag t T hr us

W e i g h t

Figure 2-13. Forces in action during flight.

Resultant

L

i f t

Relative Wind

Drag

Figure 2-12. In-flight ice f ormation adds weight, increas es drag and reduces lift.

2.10

Figure 2-14. Resultant of lift and drag. Module 08 - Basic Aerodynamics

GENERATION OF LIFT AND DRAG As stated, lift is the force that pushes the aircraft upwards. Te Angle of Attack (AOA) is the angle between the relative wind and the chord line of the wing. Within limits, lift can be increased by increasing the angle of attack, wing area, velocity, density of the air, or by changing the shape or size of the airfoil. When the force of lift on an aircraft’s wing equals the force of gravity, the aircraft maintains level flight.

consequence of the creation of lift. It is caused by the downwash at the trailing edge of the wing meeting the air that flows underneath the wing and the general movement of the vortices created by this towards the wingtip where wingtip vortices are created. Te greater the lift, the greater the pressure differential between these two flows of air which increases the induced drag. Since lift is able to be increased by increasing angle of attack, so too is induced drag. A discussion of the various types of drag and their production occurs after an examination of angle of attack.

Drag is the force that opposes the thrust created to move the aircraft forward. Induced drag is an inevitable

LIFT AND DRAG COEFFICIENTS Aerodynamicists calculate a Lift Coefficient (C �) to model all of the complex variables that contribute to the generation of lift. It incorporates the shape and area of the airfoil, the angle of attack, and various flow conditions such as air density and velocity. In short, the lift coefficient is a ratio between lift pressure and dynamic pressure and is a function of the shape of the

wing and angle of attac k. A Drag Coeff icient (C �) can also be calculated. Similarly, the drag coefficient incorporates the complex variables that contribute to the formation of drag. It is the ratio of d rag pressure to dynamic pressure. Te drag coefficient increases with the angle of attack and includes all types of drag as discussed in the section on drag below.

LIFT/DRAG RATIO Drag is the price paid to obtain lift.  he Lift to Drag ratio (L/D) is the amount of lift generated by a wing or airfoil compared to its drag. A ratio of L/D indicates airfoil efficiency. Aircraft with higher L/D ratios are more efficient than those with lower L/D

ratios. In unaccelerated flight with the lift and drag data steady, the proportions of the coefficient of lift (C�) and Coefficient of Drag (C �) can be calculated for specific AOA. ( Figure 2-15 )

CL

.2000

1.8

18

1.6

16

1.4

14

1.2

12

1.0

10

.0800

0.8

8

.0600

0.6

6

0.4

4

0.2

2

0

0

.1800

CL

.1600 )

L/ D MAX

D

C ( g a r D f o t n e i c fi f e o C

MAX

.1400 .1200

CL

.1000

.0400

g a r d / t f i L

Stall .0200 0

CD 0°

2°

4°

6°

8°

10°

12°

14°

16°

18°

20°

22°

Critical Angle of Attack

Figure 2-15. Lift Coefficients at various angles of attack. Module 08 - Basic Aerodynamics

2.11


Te L/D ratio is determined by dividing the C � by the C�, which is the same as dividing the lift equation by the drag equation. All terms except coefficients cancel out. L = Lift in pounds D = Drag Where L is the lif t force in pounds, C � is the lift coefficient, ρ is density expressed in slugs per cubic feet, V is velocity in feet per second, q is dynamic pressure per square feet, and S is the wing area in square feet. C � = Ratio of drag pressure to dynamic pressure. ypically at low angles of attack, the drag coefficient is low and small changes in angle of attack create only slight changes in t he drag coefficient. At high angles of attack, small changes in the angle of attack cause significant changes in drag.

Te above formulas represent the Coefficient of Lift (C�) and the Coefficient of Drag (C�) respectively. Te shape of an airfoil and other lift producing devices (e.g., flaps) effect the production of lift and alter with changes in the AOA. Te lift/drag ratio is used to express the relation between lift and drag and is determined by dividing the lift coefficient by the drag coefficient, C �/C�. Notice in Figure 2-15 that the lift curve (red) reaches its maximum for this particular wing section at 20° AOA, and then rapidly decreases. 15° AOA is therefore the stalling angle. Te drag curve (yellow) increases very rapidly from 14° AOA and completely overcomes the lift curve at 21° AOA. Te lift/drag ratio (green) reaches its maximum at 6° AOA, meaning that at this angle, the most lift is obtained for the least amount of drag. Note that the maximum Lift /Drag ratio (L/D �� ) occurs at one specific C � and AOA. If the aircraft is operated in steady flight at L/D �� , the total drag is at a minimum. Any AOA lower or higher than that for L/ D�� reduces the L/D and consequently increases the total drag for a given aircraft’s lift.

L = C . ρ . V2 . S 2 D = C . ρ . V2 . S 2

ANGLE OF ATTACK (AOA) Before beginning the discussion on AOA and its effect on airfoils, first consider the terms chord and Center of Pressure (CP) as illustrated in Figure 2-16.

incidence that was illustrated in Figure 2-8 , which is the angle between the chord line of the wing and the longitudinal axis of the aircraft.

As stated, the chord of an airfoil or wing section is an imaginary straight line that passes through the section from the leading edge to the trailing edge, as shown in Figure 2-16 . Te chord line provides one side of an angle that ultimately forms the angle of attack. Te other side of the angle is formed by a line indicating the direction of the relative airstream. Tus, the angle of attack or AOA is defined as the angle between the chord line of the wing and the d irection of the relative wi nd. his is not to be con fu sed with the angle of

On each part of an airfoil or wing surface, a smal l force is present. Tis force is of a different magnitude and direction from any forces acting on other areas forwa rd or rearward from this point. It is possible to add all of these small forces. hat sum is called the "resultant force" (lif t). his resultant force has mag nitude, direction, and location, and can be represented as a vector, as shown in Figure 2-16.

Angle of attack Resultant lift C h o rd l i n e

Relative airstream

Lift

C ho r d

Center of pressure

Figure 2-16. Airflow over a wing section.

2.12

Drag

he point of intersection of the resultant force line with the chord line of the airfoil is called the center of pressure (CP). he CP moves along the airfoil chord as the AOA changes. hroughout most of the flight range, the CP moves forward with increasing AOA and rearward as the AOA decreases. Te effect of increasing AOA on the CP is shown in Figure 2-17 . he AOA changes as the aircraft’s attitude changes. Since the AOA has a great deal to do with determining Module 08 - Basic Aerodynamics

lift, it is given primary consideration when designing airfoils. In a properly designed airfoil, the lift increases as the AOA is increased.

A Angle of attack = 0° Resultant Negative pressure pattern Relative airstream

When the AOA is increased gradually toward a positive AOA, the lift component increases rapidly up to a certain point and then suddenly begins to drop off. During this action the drag component increases slowly at first, then rapidly as lift begins to drop off. When the AOA increases to the angle of maximum lift, the burble point is reached. Tis is known as the critical angle . When the critical angle is reached, the air ceases to flow smoothly over the top surface of the airfoil and begins to burble or eddy. Tis means that air breaks away from the upper camber line of the wing. What was formerly the area of decreased pressure is now filled by this burbling air. When this occurs, the amount of lift drops and drag becomes excessive. Te force of gravity exerts itself, and the nose of the aircraft drops. Tis is a stall . Tus, the burble point is the stalling angle. As previously seen, the distribution of the pressure forces over the airfoil varies with the AOA. Te application of the resultant force, or CP, varies correspondingly. As this angle increases, the CP moves forward; as the angle decreases, the CP moves back. Te unstable travel of the CP is characteristic of almost all airfoils. Te efficiency of a wing is measured in terms of the lift to drag ratio (L/D). his ratio varies with the AOA but reaches a definite maximum value for a particular AOA. At this angle, the wing has reached its maximum efficiency. he shape of the airfoil is the factor that determines the AOA at which the wing is most efficient; it also determines the degree of efficiency. Research has shown that the most efficient airfoils for general use have the maximum thickness occurring about one-third of the way back from the leading edge of the wing.


Positive pressure


Center of pressure

B Angle of attack = 6° Resultant

Negative pressure pattern moves forward

Relative airstream

Positive pressure

C Angle of attack = 12° Resultant

Relative airstream

Center of pressure moves forward

Positive pressure

D Angle of attack = 18° Wing completely stalled

Positive pressure

Figure 2-17. Effect on increasing angle of attack.

2.13

DRAG here are many di fferent types of dr ag. he mos t common are parasite drag , induced drag and wave drag . Additionally, there are three types of parasite drag: 1. Form drag which results from the aerodynamic resistance to motion due to the shape of the aircraft. 2. Skin friction drag which is related to the smoothness (or roughness) of the aircraft surfaces. 3. Interference drag which occurs where surfaces with different flow characteristics meet (e.g. wing and fuselage). Briefly, induced drag is a secondary effect of the production of lift and wave drag comes into play when shock waves develop close to the surface of the aircraft during transonic or supersonic flight. In the following paragraphs, each of these types of drag will be explained in more detail. PARASITE DRAG

Parasite (parasitic) Drag (DP) is defined as all drag that is not associated with the production of lift. Parasite drag is caused by moving a solid object through a fluid medium. In aerodynamics, the fluid medium concerned is the atmosphere. Te principal components of parasite drag are form drag, friction drag and interference drag. FORM DRAG

Form drag, also known as pressure drag or profile drag, is caused by the separation of the boundary layer from a surface and the wake created by that separation. It is primarily dependent upon the shape of the object.

Te remaining examples in Figure 2-18 show how, as streamlining is increased, form drag is decreased. FRICTION DRAG

Friction drag , also known as skin friction drag, is

caused by the friction of a fluid against the surface of an object that is moving through it. It is directly proportional to the area of the surface in contact with the fluid and increases with the square of the velocity. In aerodynamics, the fluid concerned is the atmosphere. Friction Drag is created in the boundary layer due to the viscosity of the air and the resulting friction against the surface of the aircraft. Te air molecules in direct contact with the aircraft surface are most affected. As the molecules flow past the surface and past each other, the viscous resistance to that flow becomes a force which retards forward motion. Te amount of friction drag that is created per square meter of surface area is relatively small. However, as the boundary layer covers much of the surface of the aircraft, friction drag can become quite significant in larger aircraft. urbulent flow creates more friction drag than laminar flow due to its greater interaction with the surface of the airplane. Rough surfaces accelerate the transition of boundary layer airflow from laminar to turbulent. Tis increases boundary layer thickness and the airflow

Te upper example in Figure 2-18 shows the relative win d ac ross a f la t pl ate re su lts in a le ad ing ed ge stagnation point at the front of the plate that contains very hi gh static pres su re. he ai rf low attempt s to maintain contact with the surface of the plate, but the streamlines are unable to follow the sharp angles which would be required to allow them to fill in behind the plate. As a result, they separate at the trailing edge of the plate leaving a low pressure wake area behind it. he pressure differential between the leading and trailing edges of the plate causes the plate to be pushed in the direction of the relative wind and retards forward motion. Tis is form drag. o reduce form drag, aircraft surfaces which are exposed to the airflow of the relative wind are streamlined. 2.14

FLAT PLATE

SPHERE

SPHERE WITH A FAIRING

SPHERE INSIDE A HOUSING

Figure 2-18. Form Drag. Module 08 - Basic Aerodynamics

disruption within the boundary layer. Tese increases result in more air molecules being affected by the movement of the aircraft and a corresponding increase in friction drag. Friction drag can be reduced by delaying the point at which laminar flow becomes turbulent. Tis can be accomplished by smoothing the exposed surfaces of the aircraft by using flush rivets on the leading edges and through painting, cleaning, wa xing, polishing or the application of surface coatings. INTERFERENCE DRAG

Interference drag is generated by the mixing of airflow streamlines between airframe components such as the wing and the fuselage, the engine pylon and the wing or, in the case of a military or other special purpose aircraft, between the airframe and attached external stores such as fuel tanks, weapons or sensor pods. ( Figure 2-19) Interference drag is generated when the airflow across one component of an aircraft is forced to mix with the airflow across an adjacent or proximal component. If one considers two parts of the aircraft that intersect at a particular point, such as the vertical and horizontal components of the empennage, it is obvious where the point of intersection occurs. Each of these two components generate high velocity (potentially transonic or even supersonic) airflow across their respective surfaces. At the intersection of the two surfaces, there is less physical space for the airflow to occupy resulting in the turbulent mixing of the two air f lows and the production of a localized shock wave. Due to this shock wave, the resulting total drag from the empennage is greater than the sum of the drag produced individually by the vertical tail and the horizontal tail surfaces. Other significant locations which generate interference drag include the wing/fuselage junction and the wing/engine pylon or fuselage/engine pylon convergence.

As seen in Figure 2-19, interference drag can be minimized by the appropriate use of fairings and fillets to ease the transition between components. Fairings and fillets use curved surfaces to soften the transition at the junction of two aircraft components. Tis, in turn, allows the airf low streamlines to meet gradually rather than abruptly and reduces the amount of interference drag that is generated. INDUCED DRAG Induced Drag is an inevitable consequence of lift and

is produced by the passage of an airfoil (e.g. wing or tailplane) through the air. Air flowing over the top of a wing tends to flow inwards because the decreased pressure over the top surface is less than the pressure outside the wing tip. Below the wing, the air flows outwards because the pressure below the wing is greater than that outside the wing tip. A direct consequence of this is that there is a continual spilling of air upwards around the wing tip phenomenon called ‘tip effect’ or ‘end effect’. One way to appreciate why a high aspect ratio for a wing is better than a low one is that with a high aspect ratio, the proportion of air which moves in this way is reduced and therefore more of it generates lift. For the wing more generally, since the streams of air from above and below the w ing which meet a long the trailing edge are flowing at an angle to each other as they meet, they combine to form vortices, which, when viewed from the rear, rotate clockwise from the left wing and counter clockwise from the right. Te tendency is for these vortices to move outwards towards the wing tip joining up as they do so. Eventually, by the time the wing tip is reached, one large wing tip vortex has formed and is shed. Most of the air flowing off the top of a wing (downwash) continues more or less horizontally towards the empennage because it is balanced by a corresponding upwash in front of the wing leading edge. In contrast, the upwards ai r movement which leads to vortex consolidation at the tip is just outside the tip whereas the corresponding downward movement is just at the extremity of the wingspan so that the net direction of airflow past the wing is downwards. Te lift created by the wing, which is by definition at right angles to the airflow, is therefore inclined slightly backwards and thus contributes to induced drag.

Figure 2-19. A wing root can cause interference drag. Module 08 - Basic Aerodynamics

2.15


Although there must always be at least some induced drag because wings have a finite thickness, designers attempt wherever possible to reduce this flow. A required wing area can be achieved using different wing span-to-chord ratios (aspect ratios). Te larger the wing aspect ratio, the less air disturbance is created at the tip. However, for most aircraft, there are both practical limits to maximum wing span for ground maneuvering and structural issues which mean that eventually, the weight penalt y to adequately strengthen a long thin wing becomes excessive. he fact that aircraft carr y most of their fuel in the wings is also a factor in wing design. ypical transport aircraft aspect ratios range between 6:1 and 10:1. Other ways to reduce induced drag and tip vortex strength in a wing design are also based upon reducing the quantity air movement upwards at the wing tip by aiming to generate relatively more of the lift away from tips. Wing taper towards the tip assists this as does wing twist. Te Boeing 767 is a example of a twisted wing. Te inner wing is set at a higher angle of attack than the outer wing and thus generates proportionately more lift whereas the tip, at a very small angle of attack generates very little. Winglets have also become popular, including both the usual up-turned versions and the Airbus 320 series two way ‘wingtip fence’ versions. Well designed winglets can prevent about 20% of the airflow spillage at the tip, and therefore 20% of the induced drag. ( Figure 2-20 ) Induced drag and its wing tip vortices are a direct consequence of the creation of lift by the wing. Since the coefficient of lift is large when the angle of attack is large, induced drag is inversely proportional to the square of the speed whereas all other drag is directly proportional to the square of the speed. Te effect of this is that induced drag is relatively unimportant at high speed in the cruise and descent where it probably represents less than 10% of total drag. In a climb, it is more important representing at least 20% of total drag. At slow speeds just after take off and in the initial climb, it is of maximum importance and may produce as much as 70% of total drag. Finally, when looking at the potential strength of wing tip vortices, theory on induced drag must be moderated by the effect of aircraft weight. Induced drag always increases with aircraft weight.

2.16

Figure 2-20. Winglets help reduce induced drag.

WAVE DRAG Wave drag is a force, or drag that retards the forward

movement of an airplane, in both supersonic and transonic flight, as a consequence of the formation of shock waves. Wave drag is caused by the formation of shock waves around the aircraft in supersonic flight or around some surfaces of the aircraft while in transonic flight. In cruise, most civil jet aircraft fly in the Mach .75 to .85 speed range. Shock waves are typically associated with supersonic aircraft, however, they also form on an aircraft traveling at less than the speed of sound. h is occurs on th e aircra ft where loca l ai rf low is accelerated to sonic speed and then decreases back to subsonic speed. A shockwave (and associated wave drag) forms at the point the ai rf low becomes subsonic. Tis is common on aircraft airfoils. As the aircraft continues to accelerate, the area of the wing experiencing supersonic flow increases. Te shockwave moves further back on the wing and becomes larger. Boundary layer separation also increases with the increase in speed and if the speed is a llowed to increase beyond the limiting Mach number, severe buffeting, Mach tuck or "upset" (loss of control) may occur. Shock waves radiate a considerable amount of energy, resulting in drag on the aircraft. his wave drag can be reduced by incorporating one or more aerodynamic design features such as wing sweep, ultra thin wings, fuselage shape, anti shock bodies and super critical airfoils.


DRAG AND AIRSPEED

Parasitic Drag increases with the square of the airspeed, while induced drag, being a function of lift, is greatest when maximum lift is being developed, usually at low speeds. Figure 2-21 shows the relationship of parasitic drag and induced drag to each other and to total drag. Tere is an airspeed at which drag is minimum, and in theory, this is the ma ximum range speed. However, flight at this speed is unstable because a small decrease in speed results in an increase in drag, and a further fall in speed. In practice, for stable flight, maximum range is achieved at a speed a little above the minimum drag speed where a small speed decrease results in a reduction in drag.


I n d u c e d D r a g

g

Total Drag

g a r D

a r D i c t i s r a P a


Minimum Drag Minimum Drag Speed

Airspeed

Figure 2-21. Different types of drag versus airspeed.

2.17

2.18


QUESTIONS

Question: 2-1 __________________ is the rate of motion in a particular direction in relation to time.

Question: 2-5 A comparison between the wingspan and the chord of a wing is known as __________________.

Question: 2-2 When fluid flowing through a tube reaches a constriction, the speed of the fluid __________________ and the pressure of the fluid __________________.

Question: 2-6 __________________ is defined as the angle between the chord line of the wing and the direction of the relative wind.

Question: 2-3 Te boundary layer is the part of the airflow that is __________________ to the surface of the aircraft.

Question: 2-7 __________________ drag is generated by the mixing of airflow streamlines between airframe components such as the wing and fuselage.

Question: 2-4 Te __________________ of a wing is the width of the wing from the leading edge apex to the trailing edge.

Question: 2-8 What type of drag continually increases as airspeed increases?


2.19

ANSWERS

Answer: 2-1 velocity.

Answer: 2-5 aspect ratio.

Answer: 2-2 increases. decreases.

Answer: 2-6 Angle of Attack (AOA).

Answer: 2-3 closest.

Answer: 2-7 Interference.

Answer: 2-4 chord.

Answer: 2-8 Parasite Drag (DP).

2.20



¦

B1

B2

2

2

Sub-Module 03

THEORY OF FLIGHT

Knowledge Requirements 8.3 - Teory of Flight Relationship between lift, weight, thrust and drag; Glide ratio; Steady state flights, performance; Teory of the turn; Influence of load factor: stall, flight envelope and structural limitations; Lift augmentation.

Level 2




3.1

T H G I L F F O Y R O E H T

RELATIONSHIP BETWEEN LIFT, WEIGHT, THRUST AND DRAG WEIGHT, CENTER OF GRAVITY, LIFT Gravity is the pulling force that tends to draw all

bodies toward the center of the earth. Te Center of Gravity (CG) may be considered as a point at which all the weight of the aircraft is concentrated. If the aircraft were supported at its exact CG, it would balance in any attitude. Note that the CG is of major importance in an aircraft, for its position has a great bearing upon stability. he location of the CG is determined by the general design of each particular aircraft. Te designers determine how far the Center of Pressure (CP) will travel. hey then fix the CG forward of the center of pressure for the corresponding flight speed in order to provide an adequate restoring moment to retain flight equilibrium. Weight has a def init e relat io nship to lift. his relationship is simple, but important in understanding the aerodynamics of flying. Lift is the upward force on the wing acting perpendicular to the relative wind. Lift is required to counteract the aircraft’s weight (which is caused by the force of gravity acting on the mass of the aircraft). Tis weight (gravity) force acts downward through the airplane’s CG. In stabilized level flight, when the lift force is equal to the weight force, the aircraft is in a state of equilibrium and neither gains nor loses altitude. If lift becomes less than weight, the aircraft loses altitude. When lift is greater than weight, the aircraft gains altitude. LIFT

Te pilot can control lift. Any time the control yoke or stick is moved fore or aft, the Angle of Attack (AOA) is changed. As the AOA increases, lift increases (all other factors being equal). When the aircraft reaches the maximum AOA, lift begins to diminish rapidly. Tis is the stalling AOA, known as C �-�� critical AOA. Examine Figure 2-15 in Sub-Module 02 and note how the C� increases until the critical AOA is reached, then decreases rapidly with any further increase in the AOA. Before proceeding further with the topic of lift and how it can be controlled, velocity must be interjected. Te shape of the wing (or rotor) cannot be effective unless it continually keeps "attacking" new air. If an aircraft is to keep flying, the lift-producing airfoil must keep moving. 3.2

In a helicopter or gyro-plane this is accomplished by the rotation of the rotor blades. For other types of aircraft such as airplanes, weight-shift control, or gliders, air must be moving across the lifting surface. his is accomplished by the forward speed of the aircraft. Lift is proportional to the square of the aircraft’s velocity. For example, an airplane traveling at 200 knots has four times the lift as the same airplane traveling at 100 knots, if the AOA and other factors remain constant. An aircraft cannot not continue to travel in level flight at a constant altitude and maintain the same AOA if the velocity is increased. he lift would increase and the aircraft would climb as a result of the increased lift force. Terefore, to maintain the lift and weight forces in balance, and to keep the a ircraft straight and level (not accelerating upward) in a state of equilibrium, as velocity is increased, lift must be decreased. his is normally accomplished by reducing the AOA by lowering the nose. Conversely, as the aircraft is slowed, the decreasing velocity requires increasing the AOA to maintain lift sufficient to maintain flight. Tere is, of course, a limit to how far the AOA can be increased, if a stall is to be avoided. All other factors being constant, for every AOA there is a corresponding airspeed required to maintain a ltitude in steady, unaccelerated flight (true only if maintaining "level flight"). Since an airfoil always stalls at the same AOA, when the weight of the aircraft is increased, lift must also be increased. Te only method of increasing lift is by increasing velocity if the AOA is held constant just short of the "critical," or stalling, AOA. Lift and drag also vary directly with the density of the air. Density is affected by several factors: pressure, temperature, and humidity. At an altitude of 18 000 feet, the density of the air has one-half the density of air at sea level. In order to maintain its lift at a higher altitude, an aircraft must fly at a greater true airspeed for any given AOA. Warm air is less dense than cool air, and moist air is less dense than dry air. Tus, on a hot humid day, an aircraft must be flown at a greater true airspeed for any given AOA than on a cool, dry day.


If the density factor is decreased and the total lift must equal the total weight to remain in fl ight, it follows that one of the other factors must be increased. Te factor usually increased is the airspeed or the AOA, because these are controlled directly by the pilot. Lift varies directly with the wing area, provided there is no change in the wing’s planform. If the wings have the same proportion and airfoil sections, a wing with a planform area of 200 square feet lifts twice as much at the same AOA as a wing with an area of 100 square feet. wo ma jo r ae ro dynamic fac to rs fro m the pi lo t’s viewp oint are li ft and veloc ity because they can be controlled readily and accurately. Of course, the pilot can also control density by adjusting the altitude and can control wing area if the aircraft happens to have flaps of the type that enlarge wing area. However, for most situations, the pilot controls lift and velocity to maneuver an aircraft. For instance, in straight-and-level flight, cruising along at a constant altitude, altitude is maintained by adjusting lif t to match the aircraf t’s velocity or cruise airspeed, while maintaining a state of equilibrium in which lift equals weight. In an approach to landing, when the pilot wishes to land as slowly as practical, it is necessary to increase lift to near maximum to maintain lift equal to the weight of the aircraft. THRUST AND DRAG

hrust has a def inite relationsh ip wit h drag. hese relationships are quite simple, but very important in understanding the aerodynamics of flying. Wing area is mea sured in square feet and includes the part blanked out by the fuselage. Wing area is adequately described as the area of the shadow cast by the wing at high noon. ests show that lift and drag forces acting on a wing are roughly proportional to the wing area. Tis means that if the wing area is doubled,

all other variables remaining the same, the lift and drag created by the wing is doubled. If the area is tripled, lift and drag are tripled. Drag must be overcome for the aircraft to move, and movement is essential to obtain li ft. o overcome drag and move the aircraft forward, another force is essential. Tis force is thrust. Trust is derived from jet propulsion or from a propeller and engine combination. Jet propulsion theory is based on Newton’s third law of motion. Te turbine engine causes a mass of air to be moved backward at high velocity causing a reaction that moves the aircraft forward. In a propeller/engine combination, the propeller is actually two or more revolving airfoils mounted on a horizontal shaft. Te motion of the blades through the air produces lift similar to the lift on the wing, but acts in a horizontal direction, pulling the aircraft forward. Before the aircraft begins to move, thrust must be exerted. Te aircraft continues to move and gain speed until thrust and drag are equal. In order to maintain a steady speed, thrust and drag must remain equal, just as lift and weight must be equal for steady, horizontal flight. If the revolutions per minute (rpm) of the engine is reduced, the thrust is lessened, and the aircraft slows down. As long as the thrust is less than the drag, the aircraft travels more and more slowly until its speed is insufficient to support it in the air. Likewise, if the rpm of the engine is increased, thrust becomes greater than drag, and the speed of the aircraft increases. As long as the thrust continues to be greater than the drag, the aircraft continues to accelerate. When drag equals thrust, the aircraft flies at a steady speed. It is worth repeating that when lift balances weight and thrust balances drag, the aircraft is in level flight neither accelerating or slowing down.

GLIDE RATIO Te glide ratio of an airplane is the distance the airplane will, with power off, travel forward in relation to the altitude it loses. For instance, if an airplane travels 10 000 feet forward while descending 1 000 feet, its glide ratio is said to be 10 to 1.


he glide rat io is af fected by all fou r funda menta l forces that act on an airplane (weight, lift, drag, and thrust). If all factors affecting the airplane are constant, the glide ratio will be constant. Although the effect of wind will not be covered in this section, it is a very prominent force acting on the gliding distance of the airplane in relationship to its movement over the 3.3


ground. With a tailwind, the airplane will glide fart her because of the higher groundspeed. Conversely, with a headwind the airplane will not glide as far because of the slower groundspeed. Variations in weight do not affect the glide angle provided the pilot uses the correct airspeed. Since it is the Lift Over Drag (L/D) ratio that determines the distance the airplane can glide, weight will not affect the distance. Te glide ratio is based only on the relationship of the aerodynamic forces acting on the airplane. he only effect weight has is to vary the time the airplane will glide. Te heavier the airplane the higher the airspeed must be to obtain the same glide ratio. For example, if two airplanes having the same L/D ratio, but different weights, start a glide from the same altitude, the heavier airplane gliding at a higher airspeed will arrive at the same touchdown point in a shorter time. Both airplanes will cover the same distance, only the lighter airplane will take a longer time.

glide or power-off approach, the pitch attitude must be adjusted as necessary to maintain a constant airspeed. he be st spe ed for the gl id e is one at wh ic h the airplane will travel the greatest forward distance for a given loss of altitude in still air. Tis best glide speed corresponds to an angle of attack resulting in the least drag on the airplane and giving the best Lift-to-Drag ratio (L/D ��). ( Figure 3-1) Any change in the gliding airspeed will result in a proportionate change in glide ratio. Any speed, other than the best glide speed, results in more drag. herefore, as the glide airspeed is reduced or increased from the optimum or best gl ide speed, the glide ratio is also changed. When descending at a speed below the best glide speed, induced drag increases. When descending at a speed above best glide speed, parasite drag increases. In either case, the rate of descent will increase. ( Figure 3-2)

Under various flight conditions, the drag factor may change through the operation of the landing gear and/or flaps. When the landing gear or the flaps are extended, drag increases and the airspeed will decrease unless the pitch attitude is lowered. As the pitch is lowered, the glidepath steepens and reduces the distance traveled. With the power off, a windmil ling propeller also creates considerable drag, thereby retarding the airplane’s forward movement.

L/D max

o i t a R g a r D o t t f i L g n i s a e r c n I

Although the propeller thrust of the airplane is normally dependent on the power output of the engine, the throttle is in the closed position during a glide so the thrust is constant. Since power is not used during a

Increasing Angle of Attack

Figure 3-1. L/DMAX.

e e d id e S p l G t t B e s F a s o o T

o l w o S T o

Figure 3-2. Best glide speed provides the greatest forward distance for a given loss of altitude.

3.4


he pilot must never attempt to "stretch" a glide by applying back-elevator pressure and reducing the airspeed below the airplane’s recommended best glide

speed. Attempts to stretch a glide will invariably result in an increase in the rate and angle of descent and may precipitate an inadvertent stall.

POLAR CURVE A polar curve is a graph which contrasts the sink rate of an aircraft with its horizontal speed. It is used mainly to illustrate performance of a glider. Knowing the best speed to fly is important in exploiting the performance of a glider. wo of the key measures of a glider’s performance are its minimum sink rate and its best glide ratio, also known as the best glide angle. Tese occur at different speeds. Knowing these speeds is important for efficient cross country flying. In still air the polar curve shows that flying at the minimum sink speed enables the pilot to stay airborne for as long as possible and to climb as quickly as possible. But at this speed, the glider will not travel as far as if it flew at the speed for the best glide. When in sinking air, the polar curve shows that best speed to fly depends on the rate that the air is descending. Te optimal speed to fly for best cross country speed may often be considerably in excess of the speed for the best glide angle to get out of the sinking air as quickly as possible. By measuring the rate of sink at various air-speeds a set of data can be accumulated and plotted on a graph. Te points can be connected by a line known as the polar curve. Each type of glider has a unique polar curve. Te

curve can be significantly degraded with debris such as bugs, dirt, and rain on the wing. Published polar curves will often be shown for a clean wing in addition to a dirty wing with bug splats represented by small pieces of tape applied to the leading edge of the wing. Te origin for a polar curve is where the air-speed is zero and the sink rate is zero. In Figure 3-3 a line has been drawn from the origin to the point with minimum sink. Te slope of the line from the origin gives the glide angle, because it is the ratio of the distance along the airspeed axis to the distance along the sink rate axis. A whole series of lines could be drawn from the origin to each of the data points, each line showing the glide angle for that speed. However, the best glide angle is the line with the least slope. In Figure 3-4 , the line has been drawn from the origin to the point representing the best glide ratio. Te air-speed and sink rate at the best glide ratio can be read off the graph. Note that the best glide ratio is shallower than the glide ratio for minimum sink. All the other lines from the origin to the various data points would be steeper than the line of the best glide angle. Consequently, the line for the best glide angle will only just graze the polar curve, e.g. it is a tangent. Airspeed Knots

Airspeed Knots

s t o n K e t a R k n i S

s t o n K e t a R k n i S

Figure 3-3. Polar curve showing glide angle for minimum sink.

Best Glide Angle

Figure 3-4. Polar curve showing glide angle for best glide.

AERODYNAMIC FORCES IN TURNS If an aircraft were viewed in straight-and-level flight from the front ( Figure 3-5 ), and if the forces acting Module 08 - Basic Aerodynamics

on the aircraft could be seen, lift and weight would be apparent: two forces. If the aircraft were in a bank it 3.5


Level Flight

Medium Banked Turn C V o e r m t i p c o a n l e n t

T o t a l L i f t

L i f t

Horizontal Component

W e i g h t

T o t a l L i f t

Centrifugal Force

W e i g h t

Steeply Banked Turn

R e s u l t a n t l o a d

C V o e r m t i p c o a n l e n t

Centrifugal Force Horizontal Component W e i g h t

R e s u l t a n t l o a d

Figure 3-5. Forces during normal coordinated turns.

would be apparent that lift did not act directly opposite to the weight, rather it now acts in the direction of the bank. A basic truth about turns: when the aircraft banks, lift acts inward toward the center of the turn, as well as upward. Newton’s First Law of Motion, the Law of Inertia, states that an object at rest or moving in a straight line remains at rest or continues to move in a straight line until acted on by some other force. An aircraft, like any moving object, requires a sideward force to make it turn. In a normal turn, this force is supplied by banking the aircraft so that lift is exerted inward, as well as upward. Te force of lift during a turn is separated into two components at right angles to each other. One component, which acts vertically and opposite to the weight (gravity), is called the "vertical component of lift." he other, which acts horizontally toward the center of the turn, is called the "horizontal component of lift," or centripetal force. Te horizontal component of lift is the force that pulls the aircraft from a straight flightpath to make it turn. Centrifugal force is the "equal and opposite reaction" of the aircraft to the change in direction and acts equal and opposite to the horizontal component of lift. Tis explains why, in a correctly executed turn, the force that turns the aircraft is not supplied by the rudder. Te rudder is used to correct any deviation between the straight track of the nose and tail of the aircraft. A good turn is one in which the nose and tail of the aircraft track along the same path. If no rudder is used in a turn, the nose of the aircraft yaws to the outside of the turn. Te rudder is used to bring the nose back in line with the relative wind. An aircraft is not steered like a boat or an automobile. In order for an aircraft to turn, it must be banked. If it 3.6

is not banked, there is no force available to cause it to deviate from a straight flightpath. Conversely, when an aircraft is banked, it turns, provided it is not slipping to the inside of the turn. Merely banking the aircraft into a turn produces no change in the total amount of lift developed. Since the lift during the bank is divided into vertical and horizontal components, the amount of lift opposing gravity and supporting the aircraft’s weight is reduced. Consequently, the aircraft loses altitude unless additional lift is created. Tis is done by increasing the AOA until the vertical component of lift is again equal to the weight. Since the vertical component of lift decreases as the bank angle increases, the AOA must be progressively increased to produce sufficient vertical lift to support the aircraft’s weight. At a given airspeed, the rate at which an aircraft turns depends upon the magnitude of the horizontal component of lift. It is found that the horizontal component of lift is proportional to the angle of bank—that is, it increases or decreases respectively as the angle of bank increases or decreases. As the angle of bank is increased, the horizontal component of lift increases, thereby increasing the Rate of urn (RO). Consequently, at any given airspeed, RO can be controlled by adjusting the angle of bank. o prov ide a ver tical component of lift suff icient to hold altitude in a level turn, an increase in the AOA is required. Since the drag of the airfoil is directly proportional to its AOA, induced drag increases as the lift is increased. Tis, in turn, causes a loss of airspeed in proportion to the angle of bank. A small angle of bank results in a small reduction in airspeed while a large Module 08 - Basic Aerodynamics

angle of bank results in a large reduction in airspeed. Additional thrust (power) must be applied to prevent a reduction in airspeed in level turns. Te required amount of additional thrust is proportional to the angle of bank. o compensate for added lift, which would result if the airspeed were increased during a turn, the AOA must be decreased, or the angle of bank increased, if a constant altitude is to be maintained. If the angle of bank is held constant and the AOA decreased, the RO decreases. In order to maintain a constant RO as the airspeed is increased, the AOA must remain constant and the angle of bank increased. An increase in airspeed results in an increase of the turn radius, and centrifugal force is directly proportional to the radius of the turn. In a correctly executed turn, the horizontal component of lift must be exactly equal and opposite to the centrifugal force. As the airspeed is increased in a constant rate level turn, the radius of the turn increases. Tis increase in the radius of turn causes an increase in the centrifugal force, which must be balanced by an increase in the horizontal component of lift, which can only be increased by increasing the angle of bank. Normal Turn

L i f t

In a slipping turn, the aircraft is not turning at the rate appropriate to the bank being used, since the aircraft is yawed toward the outside of the turning flightpath. Te aircraft is banked too much for the RO, so the horizontal lift component is greater than the centrifugal force. ( Figure 3- 6) Equilibrium between the horizontal lift component and centrifugal force is reestablished by either decreasing the bank, increasing the RO, or a combination of the two changes. A skidding turn results from an excess of centrifugal force over the horizontal lift component, pulling the aircraft toward the outside of the turn. Te RO is too great for the angle of bank. Correction of a skidding turn thus involves a reduction in the RO, an increase in bank, or a combination of the two changes. o maintain a given RO, the angle of bank must be varied with the airspeed. his becomes particularly important in high-speed aircraft. For instance, at 400 miles per hour (mph), an aircraft must be banked approximately 44° to execute a standard-rate turn (3° per second). At this angle of bank, only about 79 percent of the lift of the aircraft comprises the vertical component of the lift. Tis causes a loss of altitude unless the AOA is increased sufficiently to compensate for the loss of vertical lift.

Slipping Turn

V e r t i c a l L i f t

L i f t

Centrifugal Force

L i f t

W e i g h t

L o a d


Centrifugal Force

Centrifugal Force Horizontal Lift

Horizontal Lift Centrifugal force equals horizontal lift

Skidding Turn


Centrifugal force less than horizontal lift

Horizontal Lift W e i g h t

L o a d

Centrifugal force greater than horizontal lift

W e i g h t

L o a d

Figure 3-6. Normal, slipping, and skidding turns.

STALLS An aircraft stall results from a rapid decrease in lift caused by the separation of airflow from the wing’s surface brought on by exceeding the critical AOA. A stall can occur at any pitch attitude or airspeed. Stalls are one of the most misunderstood areas of aerodynamics because pilots often believe an airfoil stops producing lift Module 08 - Basic Aerodynamics

when it stalls. In a stall, the wing does not totally stop producing lift. Rather, it can not generate adequate lift to sustain level flight. Since the C� increases with an increase in AOA, at some point the C � peaks and then begins to drop off. 3.7


Tis peak is called the C �-��. Te amount of lift the wing produces drops dramatically after exceeding the C�-�� or critical AOA, but as stated above, it does not completely stop producing lift. In most straight-wing aircraft, the wing is designed to stall the wing root first. Te wing root reaches its critical AOA first making the stall progress outward toward the wingtip. By having the wing root stall first, aileron effectiveness is maintained at the wingtips, maintaining controllability of the aircraft. Various design methods are used to achieve the stalling of the wing root first. In one design, the wing is "twisted" to a higher AOA at the wing root. Installing stall strips on the first 20–25 percent of the wing’s leading edge is another method to introduce a stall prematurely. he wing never completely stops producing lift in a stalled condition. If it did, the aircraft would fall to the Earth. Most training aircraft are designed for the nose of the aircraft to drop during a stall, reducing the AOA and "unstalling" the wing. Te "nose-down" tendency is due to the CL being aft of the CG. he CG range is very important when it comes to stall recovery characteristics. If an aircraft is allowed to be operated outside of the CG, the pilot may have difficulty recovering from a stall. Te most critical CG violation would occur when operating with a CG which exceeds the rear limit. In this situation, a pilot may not be able to generate sufficient force with the elevator to counteract the excess weight aft of the CG. Without the ability to decrease the AOA, the aircraft continues in a stalled condition until it contacts the ground.

maintaining altitude. Te lower the airspeed becomes, the more the AOA must be increased. Eventually, an AOA is reached which results in the wing not producing enough lift to support the aircraft which starts settling. If the airspeed is reduced further, the aircraft stalls, since the AOA has exceeded the critical angle and the airflow over the wing is disrupted. Low speed is not necessary to produce a stall. he wing can be brought into an excessive AOA at any speed. For example, an aircraft is in a dive with an airspeed of 100 knots when the pilot pulls bac k sharply on the elevator control. ( Figure 3-7 ) Gravity and centrifugal force prevent an immediate alteration of the flightpath, but the aircraft’s AOA changes abruptly from quite low to very high. Since the flightpath of the aircraft in relation to the oncoming air determines the direction of the relative wind, the AOA is suddenly increased, and the aircraft would reach the stalling angle at a speed much greater than the normal stall speed. he sta ll ing speed of an aircraft is also higher in a level turn than in straight-and-level flight. ( Fig ure 3-8 ) Centrifugal force is added to the aircraft’s weight and the wing must produce sufficient additional lift to counterbalance the load imposed by the combination of centrifugal force and weight. In a turn, the necessary additional lift is acquired by applying back pressure to the elevator control. Tis increases the wing’s AOA, and results in increased lift. Te AOA must increase as the

Te stalling speed of a particular aircraft is not a fixed value for all flight situations, but a given aircraft always stalls at the same AOA regardless of airspeed, weight, load factor, or density altitude. Each aircraft has a particular AOA where the airflow separates from the upper surface of the wing and the stall occurs. his critical AOA varies from 16° to 20° depending on the aircraft’s design. But each aircraft has only one specific AOA where the stall occurs. here are three f light situations in which the critical AOA can be exceeded: low speed, high speed, and turning. Te aircraft can be stalled in straight-and-level flight by flying too slowly. As the airspeed decreases, the AOA must be increased to retain the lift required for 3.8

L C F W

L L

W

F C

W

C F

W C F

Figure 3-7. Forces exert ed when pulling out of a dive. Module 08 - Basic Aerodynamics

d e e p S l l a t S n i e s a e r c n I t n e c r e P

100 90 e s a e r

80

c

n I

60

d e

e

p S

l l t a

40

S

a F d a o

L

20 0

r

o t

c

0°

10°

20°

30°

40°

50°

60°

70°

80°

13 12 11 10 9 8 7 6 5 4 3 2 1 90°

” G “ r o r o t c a F d a o L

Bank Angle

CL is normally located aft of the CG. Although this makes the aircraft inherently nose-heavy, downwash on the horizontal stabilizer counteracts this condition. At the point of stall, when the upward force of the wing’s lift and the downward tail force cease, an unbalanced condition exists. Tis allows the aircraft to pitch down abruptly, rotating about its CG. During this nose-down attitude, the AOA decreases and the airspeed again increases. Te smooth flow of air over the wing begins again, lift returns, and the aircraft is again flying. Considerable altitude may be lost before this cycle is complete.

Figure 3-8. Increase in stall speed and load factor when banking.

bank angle increases to counteract the increasing load caused by centrifugal force. If at any time during a turn the AOA becomes excessive, the aircraft stalls. At this point, the action of the aircraft during a stall should be examined. o balance the aircraft aerodynamically, the

As stated in Sub-Module 02, airfoil shape and degradation of that shape must also be considered in a discussion of stalls. Combined with the increased drag and reduced lift generation due to the accumulation of ice, snow or frost on the aircraft lifting surfaces, a stall may occur at a lower angle of attack than normal or at a higher speed.

FLIGHT ENVELOPE AND STRUCTURAL LIMITATIONS A flight envelope, performance envelope or service envelope refers to capabilities and limitations of a particular aircraft design package. In particular, performance airspeeds and load factors at different altitudes are considered. It is important to maintain flight "within the envelope" or a structural failure could occur. LOAD FACTORS In aerodynamics, load factor is the ratio of the maximum

load an aircraft can sustain to the gross weight of the aircraft. Te load factor is measured in Gs (acceleration of gravity), a unit of force equal to the force exerted by gravity on a body at rest and indicates the force to which a body is subjected when it is accelerated. Any force applied to an aircraft to deflect its flight from a straight line produces a stress on its structure, and the amount of this force is the load factor. For example, a load factor of 3 means the total load on an aircraft’s structure is three times its gross weight. Since load factors are expressed in terms of Gs, a load factor of 3 may be spoken of as 3 Gs, or a load factor of 4 as 4 Gs. If an aircraft is pulled up from a dive, subjecting the pilot to 3 Gs, he or she would be pressed down into the seat with a force equal to three times his or her weight. Since modern aircraft operate at significantly higher speeds Module 08 - Basic Aerodynamics

than older aircraft, increasing the magnitude of the load factor, this effect has become a primary consideration in the design of the structure of all aircraft. With th e st ruc tura l de sign of ai rc ra ft pl an ned to withstand only a certain amount of overload, a knowledge of load factors has become essential for all pilots. Load factors are important for two reasons: 1. It is possible for a pilot to impose a dangerous overload on the aircraft structures. 2. An increased load factor increases the stalling speed and makes stalls possible at seemingly safe flight speeds. LOAD FACTORS IN AIRCRAFT DESIGN Te answer to the question "How strong should an aircraft be? " is determined largely by the use to which the

aircraft is subjected. Tis is a difficult problem because the maximum possible loads are much too high for use in efficient design. It is true that any pilot can make a very hard landing or an extremely sharp pull up from a dive, which would result in abnormal loads. However, such extremely abnormal loads must be dismissed somewhat if aircraft are built that take off quickly, land slowly, and carry worthwhile payloads. 3.9


Te problem of load factors in aircraft design becomes how to determine the highest load factors that can be expected in normal operation under various operational situations. hese load factors are called "limit load factors." For reasons of safety, it is required that the aircraft be designed to withstand these load factors without any structural damage. A lthough certification requirements typically require the aircraft structure be capable of supporting one and one-half times these limit load factors without failure, it is accepted that parts of the aircraft may bend or twist under these loads and that some structural damage may occur. Tis 1.5 load limit factor is called the "factor of safety" and provides, to some extent, for loads higher than those expected under normal and reasonable operation. Tis strength reserve is not something which pilots should willfully abuse; rather, it is there for protection when encountering unexpected conditions. Te above considerations apply to all loading conditions, whether they be due to gusts, maneuvers, or landings. he gust load factor requirements now in effect are substantially the same as those that have been in existence for years. Hundreds of thousands of operational hours have proven them adequate for safety. Since the pilot has little control over gust load factors (except to reduce the aircraft’s speed when rough air is encountered), the gust loading requirements are substantially the same for most general aviation ty pe aircraft regardless of their operational use. Generally, the gust load factors control the design of aircraft which are intended for strictly nonacrobatic usage. An entirely different situation exists in aircraft design with maneuvering load factors. It is necessary to discuss this matter separately with respect to: (1) aircraft designed in accordance with the category system (e.g., normal, utility, acrobatic); and (2) older designs built according to requirements which did not provide for operational categories.

Aircraft designed under the category system are readily identified by a placard in the flight deck , which states the operational category (or categories) in which the aircraft is certificated. he maximum safe load factors (limit load factors) specified for aircraft in the various categories are: Typical Category Limit Load Factors

Normal Utility (mild acrobatics, including spins) Acrobatic

3.8 to – 1.52 4.4 to – 1.76 6.0 to – 3.00

�For aircraft with gross weight of more than 4 000 pounds, the limit load factor is reduced. o the limit loads given above, a safety factor of 50 percent is added. Also note that certification standards in various countries of manufacture may differ.

Tere is an upward graduation in load factor with the increasing severity of maneuvers. Te category system provides for maximum utility of an aircraft. If normal operation alone is intended, the required load factor (and consequently the weight of the aircraft) is less than if the aircraft is to be employed in training or acrobatic maneuvers as they result in higher maneuvering loads. Aircraft that do not have the category placard are designs that were constructed under earlier engineering requirements in which no operational restrictions were specifically given to the pilots. For aircraft of this type (up to weights of about 4 000 pounds), the required strength is comparable to present-day utility category aircraft, and the same types of operation are permissible. For aircraft of this type over 4 000 pounds, the load factors decrease with weight. Tese aircraft should be regarded as being comparable to the normal category aircraft designed under the category system, and they should be operated accordingly.

HIGH SPEED FLIGHT SUBSONIC VS SUPERSONIC FLOW In subsonic aerodynamics, the theory of lift is based upon

the forces generated on a body and a moving gas (air) in which it is immersed. At speeds of approximately 260 knots, air can be considered incompressible in that, 3.10

at a fixed altitude, its density remains nearly constant while its pressure varies. Under this assumption, air acts the same as water and is classified as a fluid. Subsonic aerodynamic theory also assumes the effects of viscosity (the property of a fluid that tends to prevent motion Module 08 - Basic Aerodynamics

of one part of the fluid with respect to another) are negligible, and classifies air as an ideal fluid, conforming to the principles of ideal-fluid aerodynamics such as continuity, Bernoulli’s principle, and circulation. In reality, air is compressible and viscous. While the effects of these properties are negligible at low speeds, compressibility effects in particular become increasingly important as speed increases. Compressibility (and to a lesser extent viscosity) is of paramount importance at speeds approaching the speed of sound. In these speed ranges, compressibility causes a change in the density of the air around an aircraft. During flight, a wing produces lift by accelerating the airflow over the upper surface. Tis accelerated air can, and does, reach sonic speeds even though the aircraft itself may be flying subsonic. At some extreme AOAs, in some aircraft, the speed of the air over the top surface of the wing may be double the aircraft’s speed. It is therefore entirely possible to have both supersonic and

subsonic airflow on an aircraft at the same time. When flow velocities reach sonic speeds at some location on an aircraft (such as the area of maximum camber on the wing), further acceleration results in the onset of compressibility effects such as shock wave formation, drag increase, buffeting, stability and control difficulties. Subsonic flow principles are invalid at all speeds above this point. ( Figure 3-9) M = 0.50


Maximum Local Velocity is Less Than Sonic Critical Mach number M = 0.72

Maximum Local Velocity Equal to Sonic n i c r s o

Normal shockwave

p e o w fl

u S

Subsonic M = 0.77

P os s ib l e S

ep e r a t i on

Figure 3-9. Wing Airflow.

SPEED RANGES he speed of sound varies wit h temperatu re. Under standard temperature conditions of 15 °C, the speed of sound at sea level is 661 knots. At 40 000 feet, where the temperature is –55 °C, the speed of sound decreases to 574 knots. In high-speed flight and/or high-altitude flight, the measurement of speed is expressed in terms of a "Mach number"—the ratio of the true airspeed of the aircraft to the speed of sound in the same atmospheric conditions. An aircraft traveling at the speed of sound is traveling at Mach 1.0. Aircraft speed regimes are defined approximately as follows: Subsonic—Mach numbers below 0.75 ransonic—Mach numbers from 0.75 to 1.20 Supersonic—Mach numbers from 1.20 to 5.00 Hypersonic—Mach numbers above 5.00 While f lights in the transonic and supersonic ranges are common occurrences for military aircraft, civilian jet aircraft normally operate in a cruise speed range of Mach 0.7 to Mach 0.90. Te speed of an aircraft in which airflow over any part of the aircraft or structure under consideration first reaches Module 08 - Basic Aerodynamics

(but does not exceed) Mach 1.0 is termed "critical Mach number" or "Mach Crit." Tus, critical Mach number is the boundary between subsonic and transonic flight and is largely dependent on the wing and airfoil design. Critical Mach number is an important point in transonic flight. When shock waves form on the aircraft, airflow separation followed by buffet and aircraft control difficulties can occur. Shock waves, buffet, and airflow separation take place above critical Mach number. A jet aircraft typically is most efficient when cruising at or near its critical Mach number. At speeds 5–10 percent above the critical Mach number, compressibility effects begin. Drag begins to rise sharply. Associated with the " drag rise " are buffet, trim and stability changes, and a decrease in control surface effectiveness. Tis is the point of "drag divergence." ( Figure 3-10 ) VMO/MMO is defined as the maximum operating limit speed. VMO is expressed in Knots Calibrated Airspeed (KCAS), while MMO is expressed in Mach number. he VMO limit is usually associated with operations at lower altitudes and deals with structural loads and flutter. he MMO limit is associated with operations at higher altitudes and is usually more 3.11

152 KCAS but the true velocity is about 287 KAS with a Mach number of 0.50. ) t n e i c fi f e o c g a r D ( D

C

Force Divergence Mach Number

C L = 0.3

Critical Mach Number

0.5 M (Mach number)

1.0

Figure 3-10. Critical Mach.

concerned with compressibility effects and flutter. At lower altitudes, structural loads and flutter are of concern; at higher altitudes, compressibility effects and flutter are of concern. Adherence to these speeds prevents structural problems due to dynamic pressure or f lutter, degradation in aircraft control response due to compressibility effects (e.g., Mach uck, aileron reversal, or buzz), and separated airflow due to shock waves resulting in loss of lift or vibration and buffet. Any of these phenomena could prevent the pilot from being able to adequately control the aircraft. For example, an early civilian jet aircraft had a VMO limit of 306 KCAS up to approximately FL 310 (on a standard day). At this altitude (FL 310), an MMO of 0.82 was approximately equal to 306 KCAS. Above this altitude, an MMO of 0.82 always equaled a KCAS less than 306 KCAS and, thus became the operating limit as you could not reach the VMO limit without first reaching the MMO limit. For example, at FL 380, an MMO of 0.82 is equal to 261 KCAS. MACH NUMBER VS AIRSPEED

It is important to understand how airspeed varies with Mach number. As an example, consider how the stall speed of a jet transport aircraft varies with an increase in altitude. he increase in altitude results in a corresponding drop in air density and outside temperature. Suppose this jet transport is in the clean configuration (gear and flaps up) and weighs 550 000 pounds. Te aircraft might stall at approximately 152 KCAS at sea level. Tis is equal to (on a standard day) a true velocity of 152 KAS and a Mach number of 0.23. At FL 380, the aircraft will still stall at approximately 3.12

Although the stalling speed has remained the same for our purposes, both the Mach number and AS have increased. With increasing altitude, the air density has decreased; this requires a faster true airspeed in order to have the same pressure sensed by the pitot tube for the same KCAS or KIAS (for our purposes, KCAS and KIAS are relatively close to each other). Te dynamic pressure the wing experiences at FL 380 at 287 KAS is the same as at sea level at 152 KAS. However, it is flying at higher Mach number. Another factor to consider is the speed of sound. A decrease in temperature in a gas results in a decrease in the speed of sound. Tus, as the aircraft climbs in altitude with outside temperature dropping, the speed of sound is dropping. At sea level, the speed of sound is approximately 661 KCAS, while at FL 380 it is 574 KCAS. hus, for our jet transport aircraft, the stall speed (in KAS) has gone from 152 at sea level to 287 at FL 380. Simultaneously, the speed of sound (in KCAS) has decreased from 661 to 574 and the Mach number has increased from 0.23 (152 KAS divided by 661 KAS) to 0.50 (287 KAS divided by 574 KAS). All the while the KCAS for stall has remained constant at 152. Tis describes what happens when the aircraft is at a constant KCAS with increasing altitude, but what happens when the pilot keeps Mach constant during the climb? In normal jet flight operations, the climb is at 250 KIAS (or higher (e.g. heavy)) to 10 000 feet and then at a specified en route climb airspeed (such as about 330 if a DC10) until reaching an altitude in the "mid-twenties" where the pilot then climbs at a constant Mach number to cruise altitude. Assuming for illustration purposes that the pilot climbs at a MMO of 0.82 from sea level up to FL 380. KCAS goes from 543 to 261. Te KIAS at each altitude would follow the same behavior and just differ by a few knots. Recall from the earlier discussion that the speed of sound is decreasing with the drop in temperature as the aircraft climbs. he Mach number is simply the ratio of the true airspeed to the speed of sound at flight conditions. Te significance of this is that at a constant Mach number climb, the KCAS (and KAS or KIAS as well) is falling off.


If the aircraft climbed high enough at this constant MMO with decreasing K IAS, KCAS, and KAS, it would begin to approach its stall sta ll speed. At some point the stall speed of the aircraft in Mach number could equal the MMO of the aircraft, and the pilot could neither slow up (without (without stalling) stal ling) nor speed up (without exceeding the max operating speed of the aircraft). Tis has been dubbed the "coffin corner."

A supersonic airstream passing through a normal shock wave experiences these changes: • Te airstream is slowed to subsonic. • Te airflow immediately immediately behind the shock wave does not change direction. • Te static pressure and density of of the airstream behind the wave is greatly increased. • Te energy of the airstream (indicated by total total pressure—dynamic plus static) is greatly reduced.

SHOCK WA WAVES VES

When an airplane air plane flies at subsonic speeds, the air ahead is "warned" of the airplane’s coming by a pressure change transmitted ahead of the airplane at the speed of sound. Because of this warning, the air begins to move aside before the airplane arrives and is prepared to let it pass easily. When the airplane’s speed reaches the speed of sound, the pressure change can no longer warnn the air ahead because the air war airplane plane is keeping k eeping up withh its own press wit pressure ure wave waves. s. Rat Rather, her, the air par partic ticles les pile up in front of the airplane causing a sharp decrease in the flow velocity directly in front of the airplane with a corresponding increase in air pressure and density. As the airplane’s speed increases beyond the speed of sound, the pressure and density of the compressed air ahead of it increase, the area of compression extending some distance ahead of the airplane. At some point in the airstream, the air particles are completely undisturbed, having had no advanced warning of the airplane’s approach, and in the next instant the same air particles are forced to undergo sudden and drastic changes in temperature, pressure, density, and velocity. Te boundary between the undisturbed air and and the the region region of compressed air is called ca lled a shock or "compression" "compression" wave. Tis same type ty pe of wave is formed formed whenever a supersonic supersonic airstream is slowed to subsonic without a change in direction, such as when the airstream is accelera accelerated ted to sonic speed over the cambered portion of a wing, and then decelerated to subsonic speed as the area of maximum camber is passed. A shock wave forms forms as a boundary between the supersonic and subsonic ranges. Whene W henever ver a shoc k wav wavee for forms ms per pen pendi dicu cu la larr to the airflow, it is termed a "normal "normal"" shock wave, and the flow immediately behind the wave is subsonic.


Shock wave formation causes an increase in drag. One of the principal effects of a shock wave is the formation of a dense high pressure region immediately behind the wave. Te instabili instability ty of the high pressure region region,, and the fact that part of the velocity energy of the airstream is converted to heat as it flows through the wave is a contributing factor in the drag increase, but the drag resulting from airflow separation is much greater. If the shock wave is strong, the boundary layer may not have sufficient kinetic energy to withstand airflow separation. Te drag incurred in the transonic region due to shock wave formation and airflow separation is known as "wave drag ." ." When speed exceeds the critical Mach number by about 10 percent, wave drag increases sharply. A considerable increase in thrust (power) is required to increase flight speed beyond this point into the supersonic range where, depending on the airfoil shape and the angle of attack, the boundary layer may reattach. Normal shock waves form on the w ing’s upper surface and form an additional area of supersonic flow and a normal shock wave on the lower surface. As flight speed approaches the speed of sound, the areas of supersonic flow enlarge and the shock waves move nearer the trailing edge. ( Figure 3-11) Associated with " drag rise " are buffet (known as Mach buffet), trim and stability changes, and a decrease in control force effectiveness. Te loss of lift due to airflow separation results in a loss of downwash, and a change chan ge in the position of the center pressure on the wing. Airflow separation produces a turbulent wake behind the wing, whichh caus whic causes es the tai taill sur surface facess to buf buffet fet (vibrate). he nose-up and nose-down pitch control provided by the horizontal tail is dependent on the downwash behind the wing. hus, an increase in downwash decreases the horizontal tail’s pitch control effectiveness since it effectively increases the angle of attack that the tail surface is seeing. 3.13


Movement of the wing CP affects the wing pitching moment. momen t. If the CP moves aft, af t, a diving div ing moment referred referred to as "Mach tuck" or "tuck under" is produced, and if it moves forward, a nose-up moment is produced. Tis is the primary reason for the development of the -tail configuration on many turbine-powered aircraft, which places the horizontal stabilizer as far as practical from the turbulence of the wings.

l o w i c F

o n r s

p e u

Normal Shockwave

S

M = 0.82

Normal Shockwave

i c s o n

p e r S u

F l o w

M = 0.95

SWEEPBACK

Most of the diffic difficulties ulties of transonic flight fli ght are associated withh shoc wit shockk wav wavee ind induced uced flow sepa rat ration. ion. Tere Terefore, fore, any means of delaying or allev iating the shock induced separation improves aerodynamic performance. One method is wing sweepback. Sweepback theory is based upon the concept that it is only the component of the airflow perpendicular to the leading edge of the w ing that affects pressure distribution and formation of shock waves. ( Figure 3-12)

Bow Wave

M = 1.05

Subsonic Airflow

Figure 3-11. Shock waves.

On a straight wing aircraft, the airflow strikes the wing leading edge at 90°, and its full fu ll impact produces pressure and lift. A wing with sweepback is struck by the same airflow at an angle smaller than 90°. Tis airflow on the swept wing has the effect of persuading the wing into believing that it is flying slower than it really is; thus the formation of shock waves is delayed. Advantages of wing sweep include an increase in critical Mach number, force divergence Mach number, and the Mach number at which drag rises peaks. In other words, sweep delays the onset of compressibility effects. Te Mach number, which produces a sharp sh arp chan change ge in drag coefficient, is termed the "force divergence". Mach number for most airfoils, usually exceeds the critical Mach number by 5 to 10 percent. At this speed, the airflow separation induced by shock wave formation can create significant variations in the drag, lift, or pitching moment coefficients. In addition to the delay of the onset of compressibility effects, sweepback reduces the magnitude in the changes of drag, lift or moment coefficients. In other words, the use of sweepback "softens" the force divergence. A disadvantage of swept wings is that they tend to stall at the wingtips rather than at the wing w ing roots. ( Figure 3-13 3-13) Tis is because b ecause the bounda boundary ry layer tends to flow span wisee towa wis toward rd the tips and to sepa separate rate near the lead leading ing edges. Because the tips of a swept wing are on the 3.14

A lo w i r r sp F is e T e e n w d s r a u p S e e A i r s p e e d M a c h 0 . 8 5

ns e d b y W i ng M a c h 0 .7 7 0

Figure 3-12. Sweepback effect. Module 08 - Basic Aerodynamics

aft part of the wing behind the center of lift (CL), a wingtip win gtip stal stalll causes c auses the center of lift l ift to move forwa forward rd on the wing, forcing the nose to rise further. he tendency for tip stall is greatest when wing sweep and taper are combined.  he st a l l si he sitt ua t io ionn ca n be ag g ra v at ated ed by a -ta i l configuration, which affords little or no pre-stall warning in the form of tail control surface buffet. ( Figure 3-14 ) Te -tail, -tail, being above the wing wake remains effective even after the wing has begun to stall, allowing the pilot to inadvertently drive the wing into a deeper stall at a much greater AOA. If the horizontal tail surfaces then become buried in the wing’s wake, the elevator may lose all effectiveness, making it impossible to reduce pitch attitude and break the stall. In the pre-stall and immediate post-stall regimes, the lift/drag qualities of a swept wing aircraft (specifically the enormous increase in drag at low speeds) can cause an increasingly descending flightpath with no change in pitch attitude, further increasing the AOA. In this situation, without reliable AOA information, a nose-down pitch attitude with an increasing airspeed is no no guarantee that recovery has been effected, and up-elevator movement at this stage may merely keep the aircraft stalled. It is a characteristic characterist ic of -tail -tail aircraft aircra ft to pitch up viciously when sta stalle lledd in ext extreme reme nose-h nose-high igh att attitud itudes, es, mak ing recovery difficult or violent. Te stick pusher inhibits this type of stall. At approximately one knot above stall speed, pre-programmed stick forces automatically move the stick forward, preventing the stall from developing. A G-limiter may also be incorporated into the system to prevent the pitch down generated by the stick pusher from imposing excessive loads on the aircraft. A "stick shaker," on the other hand provides stall warning when the airspeed is five to seven percent above stall speed.

Intitial Stall Area


Figure 3-13. 3-13. Wingtip s tall.

Prestall

Stalled

Figure 3-14. T-tail stall. Module 08 - Basic Aerodynamics

3.15

LIFT AUGMENT AUGMENTA ATION Secondary flight controls are fitted to most aircraft to increase lift when needed during takeoff, landing and low-level lowlevel slow flight. fl ight. hese are typically airfoils that are fitted to the leading edge and or trailing edge of the wings and are a re controlled from the flight deck. FLAPS are the most common high-lift devices used Flaps are

on aircraft. Tese surfaces, su rfaces, which are attached to the trailing edge edg e of the wing, increase both both lift and induced drag for any given AOA. Flaps allow a compromise between high cruising speed and low landing speed, because they may be extended when needed, a nd retracted into the wing’s structure when not needed. here h ere ar aree fou fourr com common mon t y pe pess of f lap laps: s: pla in in,, spl it, slotted, and Fowler flaps. ( Figure 3-15 ) h e pl he plai ainn f la lapp is th thee si simpl mples estt of th thee fou r t y pe pes. s. It increases the airfoil ca mber mber,, resulting in a significant increase in the coefficient of lift lif t (C�) at a given AO AOA. A. At the same time, it greatly increases drag and moves the center of pressure (CP) aft on the airfoil, resulting in a nose-down pitching momen moment. t. Te split flap is deflected from the lower surface sur face of the airfoil and produces a slightly greater increase in lift than the plain flap. More drag is created because of the turbulent air pattern produced behind the airfoil. When fully extended, both plain and split flaps produce high drag with little additional lift.

Basic Section

Plain Flap

Split Flap

Slotted Flap

Fowler Flap

Slotted Fowler Flap

Te most popu popular lar flap on air aircra craft ft today is the slott slotted ed flap. Variations of this design are used for small aircraft, as well as for large ones. Slotted flaps increase the lift coefficient significantly more than plain or split flaps. On small aircraft, the hinge is located below the lower surface of the flap, and when the flap is lowered, a duct forms between the flap well in the wing and the leading edge of the flap. When the slotted flap is lowered, high energy air from the lower surface is ducted to the flap’s upper surface. Te high energy air from the slot accelerates the upper surface boundary layer and delays airflow separation, providing a higher C �. Tus, the slotted flap produces much greater increases in maximum coefficient of lift (C�-��) than the plain or split flap. While there are many types of slotted flaps, large aircraft often have double and even triple3.16

Figure 3-15. Four common types of flaps.

slotted flaps. hese allow the maximum increase in drag without the airflow over the flaps separating and destroying the lift they produce. Fowler flaps are a type of slotted flap. Tis flap design not only changes t he camber of the wing, it also increases the wing area. Instead of rotating down on a hinge, it slides backwards on tracks. In the first portion of its extension, it increases the drag very little, but Module 08 - Basic Aerodynamics

increases the lift a great deal as it increases both the area and camber. As the extension continues, the flap deflects downward. During the last portion of its travel, the flap increases the drag with little additional increase in lift.

Fixed Slot

LEADING EDGE DEVICES

High-lift devices also can be applied to the leading edge of the airfoil. Te most common types are fixed slots, movable slats, leading edge flaps, and cuffs. ( Figure 3-16) Fixed slots direct airflow to the upper wing surface and delay airflow separation at higher angles of attack. Te slot does not increase the wing camber, but allows a higher maximum C � because the stall is delayed until the wing reaches a greater AOA. Movable slats consist of leading edge segments, which move on tracks. At low angles of attack, each slat is held flush against the wing’s leading edge by the high pressure that forms at the wing’s leading edge. As the AOA increases, the high-pressure area moves aft below the lower surface of the wing, allowing the slats to move forward. Some slats, however, are pilot operated and can be deployed at any AOA. Opening a slat allows the air below the wing to flow over the wing’s upper surface, delaying airflow separation. Leading edge flaps, like trailing edge flaps, are used to increase both C �-�� and the camber of the wings. Tis ty pe of leading edge device is frequently used in conjunction with trailing edge flaps and can reduce the nose-down pitching movement produced by the latter. As is true with trailing edge flaps, a small increment of leading edge flaps increases lift to a much greater extent than drag. As greater amounts of flaps are extended, drag increases at a greater rate than lift. Leading edge cuffs, like leading edge flaps and trailing edge f laps are used to increase both C �-�� and the camber of the wings. Unlike leading edge flaps and trailing edge flaps, leading edge cuffs are fixed aerodynamic devices. In most cases leading edge cuffs extend the leading edge down and forward. Tis causes the airflow to attach better to the upper surface of the wing at higher angles of attack, thus lowering an aircraft’s stall speed. Te fixed nature of leading edge cuffs extracts a penalty in maximum cruise airspeed, but recent advances in design and technology have reduced this penalty.


Movable Slot


Leading Edge Flap

Leading Edge Cuff

Figure 3-16. Leading edge high lift devices.

Figure 3-17. A winglet reduces induced drag caused by air spilling off of the wingtip.

3.17

FIXED AIRFLOW DEVICES A winglet is an obvious vertical upturn of the wing’s tip resembling a vertical stabilizer. It is an aerodynamic device designed to reduce the drag created by wing tip vortices in flight. Usually made from aluminum or composite materials, winglets can be designed to optimize performance at a desired speed. ( Figure 3-17 ) Vortex generators are smal l airfoil sections usually attached to the upper surface of a wing. ( Figure 3-18 )

Tey are designed to promote positive laminar airflow over the wing and control surfaces. Usually made of aluminum and installed in a span-wise line or lines, the vortices created by these devices swirl downward assisting maintenance of the boundary layer of air flowing over the wing. Tey can also be found on the fuselage and empennage. A chord-wise barrier on the upper surface of the wing, called a stall fence , sometimes called a strake, is used to halt the span-wise flow of air. ( Figure 3-19)

During low speed f light, this can maintain proper chord-wise airflow reducing the tendency for the wing to stall. Usually made of aluminum, the fence is a fixed structure most common on swept wings, which have a natural span-wise tending boundary air flow. Often, a gap can exist between the stationary trailing edge of a wing or stabilizer and the movable control surface(s). At high angles of attack, high pressure air from the lower wing surface can be disrupted at this gap. Te result can be turbulent airflow, which increases drag. Tere is also a tendency for some lower wing boundary air to enter the gap and disrupt the upper wing surface airflow, which in turn reduces lift and control surface responsiveness. Te use of gap seals is common to promote smooth airflow in these gap areas. Gap seals can be made of a wide variety of materials ranging from aluminum and impregnated fabric to foam and plastic. Figure 3-20 shows some gap seals installed on various aircraft.

Stall Fence

Figure 3-18. Vortex generators.

Figure 3-19. A stall fence halts the flow of span-wise air.

Aileron Gap Seal

Tab Gap Seal

Figure 3-20. Gap seals promote the smooth flow of air over gaps between fixed and movable surfaces.

3.18


QUESTIONS

Question: 3-1 __________________ is the upward force on the wing acting perpendicular to the relative wind.

Question: 3-5 Te ratio of true airspeed to the speed of sound is known as __________________.

Question: 3-2 o overcome __________________ and move the aircraft forward, thrust is needed.

Question: 3-6 Te effect of a sweptback wing is a delay in the onset of __________________ effects.

Question: 3-3 An object at rest or moving in a straight line remains at rest or continues to move in a straight line until __________________.

Question: 3-7 Te most common high-lift device fitted to the aft edge of the wing which increases lift and drag is known as a __________________.

Question: 3-4 Te ratio of the maximum load an aircraft can sustain to the gross weight of the aircraft is called __________________.

Question: 3-8 A __________________ is a small airfoil section attached to the upper surface of a wing that promotes positive laminar airflow over the wing and control surfaces.


3.19

ANSWERS

Answer: 3-1 lift.

Answer: 3-5 mach number.

Answer: 3-2 drag.

Answer: 3-6 compressibility.

Answer: 3-3 acted on by some other force.

Answer: 3-7 flap.

Answer: 3-4 load factor.

Answer: 3-8 vortex generator.

3.20



¦

B1

B2

Sub-Module 04

FLIGHT STABILITY AND DYNAMICS

Knowledge Requirements 8.4 - Flight Stability and Dynamics Longitudinal, lateral and directional stability (active and passive).

2

2

Level 2




4.1

D N A Y S T I C L I I M B A A T N S Y D T H G I L F

THE AXES OF AN AIRCRAFT Whenever an aircraft changes its attitude in flight, it must turn about one or more of three axes. Figure 4-1 shows the three axes, which are imaginary lines passing through the center of the aircraft.

Lateral Axis

Te axes of an aircraft can be considered as imaginary axles around which the aircraft turns like a wheel. At the center, where all three axes intersect, each is perpendicular to the other two. Te axis that extends lengthwise through the fuselage from the nose to the

Elevator

Aileron Rudder

Aileron

Longitudinal Axis

Vertical Axis

A

Banking (roll) control affected by aileron movement

Normal Altitude

Longitudinal Axis

B

Climb and dive (pitch) control affected by elevator movement

C

Directional (yaw) control affected by rudder movement Normal Altitude

Lateral Axis

Vertical Axis Normal Altitude

Figure 4-1. Motion of an aircraft about its axes.

4.2


tail is called the longitudinal axis. Te axis that extends crosswise from wing tip to wing tip is the lateral, or pitch, axis. Te axis that passes through the center, from top to bottom, is called the vertical, or yaw, axis. Roll, pitch, and yaw are controlled by three control surfaces.

Roll is produced by the ailerons, which are located at the trailing edges of the wings. Pitch is affected by the elevators, the rear portion of the horizontal tail assembly. Yaw is controlled by the rudder, the rear portion of the vertical tail assembly.

STABILITY AND CONTROL An aircraft must have sufficient stability to maintain a uniform f lightpath and recover from the various upsetting forces. Also, to achieve the best performance, the aircraft must have the proper response to the movement of the controls. Control is the pilot action of moving the flight controls, providing the aerodynamic force that induces the aircraft to follow a desired flightpath. When an aircraft is said to be controllable, it means that the aircraft responds easily and promptly to movement of the controls. Different control surfaces are used to control the aircraft about each of the three axes. Moving the control surfaces on an aircraft changes the airflow over the aircraft’s surface. Tis, in turn, creates changes in the balance of forces acting to keep the aircraft flying straight and level. Tree terms that appear in any discussion of stability and control are: stability, maneuverability, and controllability. Stability is the characteristic of an aircraft that tends to cause it to fly (hands off) in a straight-and-level flightpath. Maneuverability is the characteristic of an aircraft to be directed along a desired flightpath and to withstand the stresses imposed. Controllability is

Positive Static Stability

the quality of the response of an aircraft to the pilot’s commands while maneuvering the aircraft. Tere are two kinds of stability, static and dynamic. STATIC STABILITY Static stability refers to the initial tendency, or direction

of movement, back to equilibrium. In aviation, it refers to the aircraft’s initial response when disturbed from a given AOA, slip, or bank. • Positive static stability —the initial tendency of the aircraft to return to the original state of equilibrium after being disturbed ( Figure 4-2) • Neutral static stability —the initial tendency of the aircraft to remain in a new condition after its equilibrium has been disturbed ( Figure 4-2) • Negative static stability —the initial tendency of the aircraft to continue away from the original state of equilibrium after being disturbed ( Figure 4-2) DYNAMIC STABILITY

Static stability has been defined as the initial tendency to return to equilibrium that the aircraft displays after being disturbed from its trimmed condition.

Neutral Static Stability

Applied Force

Applied Force

G C

Negative Static Stability

Applied Force

G C

G C G C

Figure 4-2. Three types of static stability. Module 08 - Basic Aerodynamics

4.3


Occasionally, the initial tendency is different or opposite from the overall tendency, so a distinction must be made between the two. Dynamic stability refers to the aircraft response over time when disturbed from a given AOA, slip, or bank. Tis type of stability also has three subtypes: ( Figure 4-3) • Positive dynamic stability—over time, the motion of the displaced object decreases in amplitude and, because it is positive, the object displaced returns toward the equilibrium state. • Neutral dynamic stability—once displaced, the displaced object neither decreases nor increases in amplitude. A worn automobile shock absorber exhibits this tendency. • Negative dynamic stability—over time, the motion of the displaced object increases and becomes more divergent. Stability in an aircraft affects two areas significantly: • Maneuverability—the quality of an aircraft that permits it to be maneuvered easily and to withstand the stresses imposed by maneuvers. It is governed by the aircraft’s weight, inertia, size and location of flight controls, structural strength, and powerplant. It too is an aircraft design characteristic. • Controllability—the capability of an aircraft to respond to the pilot’s control, especially with regard to flightpath and attitude. It is the quality of the aircraft’s response to the pilot’s control application when maneuvering the aircraft, regardless of its stability characteristics.

LONGITUDINAL STABILITY (PITCHING)

In designing an aircraft, a great deal of effort is spent in developing the desired degree of stability around all three axes. But longitudinal stability about the lateral axis is considered to be the most affected by certain variables in various flight conditions. Longitudinal stability is the quality that makes an aircraft

stable about its lateral ax is. It involves the pitching motion as the aircraft’s nose moves up and down in flight. A longitudinally unstable aircraft has a tendency to dive or climb progressively into a very steep dive or climb, or even a stall. Tus, an aircraft with longitudinal instability becomes difficult and sometimes dangerous to fly. Static longitudinal stability or instability in an aircraft, is dependent upon three factors: 1. Location of the wing with respect to the CG. 2. Location of the horizontal tail surfaces with respect to the CG. 3. Area or size of the tail surfaces. In analyzing stability, it should be recalled that a body free to rotate always turns about its CG. 4. o obtain static longitudinal stability, the relation of the wing and tail moments must be such that, if the moments are initially balanced and the aircraft is suddenly nose up, the wing moments and tail moments change so that the sum of their forces provides an unbalanced but restoring moment which, in turn, brings the nose down again. Similarly, if the aircraft is nose down, the resulting change in moments brings the nose back up.

Damped Oscillation Undamped Oscillation Positive Static (positive dynamic)

t n e m e c a l p s i D

Divergent Oscillation

Time

Positive Static (neutral dynamic)

Positive Static (negative dynamic)

Figure 4-3. Damped versus undamped stability.

4.4


Te Center of Lift (CL) in most asymmetrical airfoils has a tendency to change its fore and aft positions with a change in the AOA. Te center of lift tends to move forward with an increase in AOA and to move aft with a decrease in AOA. his means that when the AOA of an airfoil is increased, the center of lift, by moving forward, tends to lift the leading edge of the wing still more. Tis tendency gives the wing an inherent quality of instability. (Note: center of lift is also known as the Center of Pressure (CP)). Figure 4- 4 shows an aircraft in straight-and-level

flight. he line CG-CL- represents the aircraft’s longitudinal axis from the CG to a point  on the horizontal stabilizer. Most aircraft are designed so that the wing’s CL is to the rear of the CG. Tis makes the aircraft "nose heavy" and requires that there be a slight downward force on the horizontal stabilizer in order to balance the aircraft and keep the nose from continually pitching downward. Compensation for this nose heaviness is provided by setting the horizontal stabilizer at a slight negative AOA. he downward force thus produced holds the tail down, counterbalancing the "heavy" nose. It is as if the line CG-CL- were a lever with an upward force at CL and two downward forces balancing each other, one a strong force at the CG point and the other, a much lesser force, at point  (downward air pressure on the stabilizer). o better visualize this physics principle: If an iron bar were suspended at point CL, with a heavy weight hanging on it at the CG, it would take downward pressure at point  to keep the "lever" in balance. Even though the horizontal stabilizer may be level when the aircraft is in level flight, there is a downwash of air from the wings. Tis downwash strikes the top of the stabilizer and produces a downward pressure, which at

a certain speed is just enough to balance the "lever." Te faster the aircraft is flying, the greater this downwash and the greater the downward force on the horizontal stabilizer (except -tails). ( Figure 4-5 ) In aircraft with fixed-position horizontal stabilizers, the aircraft manufacturer sets the stabilizer at an angle that provides the best stability (or balance) during flight at the design cruising speed and power setting. If the aircraft’s speed decreases, the speed of the airflow over the wing is decreased. As a result of this decreased flow of air over the wing, the downwash is reduced, causing a lesser downward force on the horizontal stabilizer. In turn, the characteristic nose heaviness is accentuated, causing the aircraft’s nose to pitch down more. ( Figure 4- 6) Tis places the aircraft in a nose-low attitude, lessening the wing’s AOA and drag and allowing the airspeed to increase. As the aircraft continues in the nose-low attitude and its speed increases, the downward force on the horizontal stabilizer is once again increased. Consequently, the tail is again pushed downward and the nose rises into a climbing attitude. As this climb continues, the airspeed again decreases, causing the downward force on the tail to decrease until the nose lowers once more. Because the aircraft is dynamically stable, the nose does not lower as far this

CG Balanced

tail load

Cruise Speed

CG Lesser downward tail load

Low Speed

CG Greater downward tail load

High Speed

Figure 4-4. Longitudinal stability. Module 08 - Basic Aerodynamics

Figure 4-5. Effect of speed on downwash.

4.5


L i f t

Thrust

CG

rust

CG W e i g h t

Below center of gravity

Normal Downwash

L i f t

Thrus t

CG W e i g h t

rust

CG

rough center of gravity

Reduced Downwash

Figure 4-6. Reduced power allows pitch down.

rust

CG

time as it did before. Te aircraft acquires enough speed in this more gradual dive to start it into another climb, but the climb is not as steep as the preceding one. After several of these diminishing oscillations, in which the nose alternately rises and lowers, the aircraft finally settles down to a speed at which the downward force on the tail exactly counteracts the tendency of the aircraft to dive. When this condition is attained, the aircraft is once again in balanced flight and continues in stabilized flight as long as this attitude and airspeed are not changed.

Above center of gravity Figure 4-7. Thrust line affects longitudinal stability.

L

i f t

Thrust

CG

Cruise Power

A similar effect is noted upon closing the throttle. Te downwash of the wings is reduced and the force at  in Fig ure 4- 4 is not enough to hold the horizontal stabilizer down. It seems as if the force at  on the lever were allowing the force of gravity to pull the nose down. Tis is a desirable characteristic because the aircraft is inherently trying to regain airspeed and reestablish the proper balance. Power or thrust can also have a destabilizing effect in that an increase of power may tend to make the nose rise. Te aircraft designer can offset this by establishing a "high thrust line" wherein the line of thrust passes above the CG. ( Figures 4-7 and 4-8 ) In this case, as power or thrust is increased a moment is produced to counteract the down load on the tail. On the other hand, a very "low thrust line" would tend to add to the nose-up effect of the horizontal tail surface.

4.6

L i f t

Thrust

CG

Idle Power L

i f t

Thrust

CG

Full Power

Figure 4-8. Power changes affect longitudinal stability.

Conclusion: with center of gravity forward of the center of lift and with an aerodynamic tail-down force, the aircraft usually tries to return to a safe flying attitude.


LATERAL STABILITY (ROLLING)

Stability about the aircraft’s longitudinal axis, which extends from the nose of the aircraft to its tail, is called lateral stability . his helps to stabilize the lateral or "rolling effect" when one wing gets lower than the wing on the opposite side of the aircraft. here are four ma in de sign fa ctors that ma ke an aircraft laterally stable: dihedral, sweepback, keel effect, and weight distribution.

SWEEPBACK

Sweepback is an addition to the dihedral that increases

the lift created when a wing drops from the level position. A sweptback wing is one in which the leading edge slopes backward. When a disturbance causes an aircraft with sweepback to slip or drop a wing, the low wing presents its leading edge at an angle that is perpendicular to the relative airflow. As a result, the low wing acquires more lift, rises, and the aircraft is restored to its original flight attitude.

DIHEDRAL

he most common procedure for producing lateral stability is to build the wings with an angle of one to three degrees above perpendicular to the longitudinal axis. Te wings on either side of the aircraft join the fuselage to form a slight V or angle called " dihedral ." Te amount of dihedral is measured by the angle made by each wing above a line parallel to the lateral axis. Dihedral involves a balance of lift created by the wings’ AOA on each side of the aircraft’s longitudinal axis. If a momentary gust of wind forces one wing to rise and the other to lower, the aircraft banks. When the aircraft is banked without turning, the tendency to sideslip or slide downward toward the lowered wing occurs. ( Figure 4-9) Since the wings have dihedral, the air strikes the lower wing at a much greater AOA than the higher wing. Te increased AOA on the lower wing creates more lift than the higher wing. Increased lift causes the lower wing to begin to rise upward. As the wings approach the level position, the AOA on both wings once again are equal, causing the rolling tendency to subside. Te effect of dihedral is to produce a rolling tendency to return the aircraft to a laterally balanced flight condition when a sideslip occurs. Te restoring force may move the low wing up too far, so that the opposite wing now goes down. If so, the process is repeated, decreasing with each lateral oscillation until a balance for wings-level flight is finally reached. Conversely, excessive dihedral has an adverse effect on lateral maneuvering qualities. Te aircraf t may be so stable laterally that it resists an intentional rolling motion. For this reason, aircraft that require fast roll or banking characteristics usually have less dihedral than those designed for less maneuverability. Module 08 - Basic Aerodynamics

Sweepback also contributes to directional stability. When tur bu lence or rudder application causes the aircraft to yaw to one side, the right wing presents a longer leading edge perpendicular to the relative airflow. Te airspeed of the right wing increases and it acquires more drag than the left wing. Te additional drag on the right wing pulls it back, turning the aircraft back to its original path. KEEL EFFECT AND WEIGHT DISTRIBUTION

An aircraf t always has the tendency to turn the longitudinal axis of the aircraft into the relative wind. Tis "weather vane" tendency is similar to the keel of a ship and exerts a steadying influence on the aircraft laterally about the longitudinal a xis. When the aircraft is disturbed and one wing dips, the fuselage weight acts like a pendulum returning the airplane to its original attitude. Laterally stable aircraft are constructed so that the greater portion of the keel area is above and behind the CG. ( Figure 4-10 )

Normal Angle of Attack

Dihedral Lateral Stability

Lesser Angle of Attack

Greater Angle of Attack

Figure 4-9. Dihedral for lateral stability.

4.7


CG

CG

CG Centerline

Area Forward of CG

CG

Area After Center of Gravity (CG)

Figure 4-10. Keel area for lateral st ability.

Tus, when the aircraft slips to one side, the combination of the aircraft’s weight and the pressure of the airflow against the upper portion of the keel area (both acting about the CG) tends to roll the aircraft back to wingslevel flight.

w

a

y

C G y

a w

R el at i v e W in

d

VERTICAL (DIRECTIONAL) STABILITY (YAWING)

Stability about the aircraft’s vertical axis (the sideways moment) is called yawing or directional stability . Yawing or directional stability is the most easily achieved stability in aircraft design. Te area of the vertical fin and the sides of the fuselage aft of the CG are the prime contributors which make the aircraft act like the well known weather vane or arrow, pointing its nose into the relative wind. In examining a weather vane, it can be seen that if exactly the same amount of surface were exposed to the wind in front of the pivot point as behind it, the forces fore and aft would be in balance and little or no directional movement would result. Consequently, it is necessary to have a greater surface aft of the pivot point than forward of it. Similarly, the aircraft designer must ensure positive directional stability by making the side surface greater aft than ahead of the CG. ( Figure 4-11 ) o provide additional positive stability to that provided by the fuselage, a vertical fin is added. Te fin acts similar to the feather on an arrow in maintaining straight flight. Like the weather vane and the arrow, the farther aft this fin is placed and the larger its size, the greater the aircraft’s directional stability. If an aircraft is flying in a straight line, and a sideward gust of air gives the aircraft a slight rotation about its vertical axis (e.g., the right), the motion is retarded and stopped by the fin because while the aircraft is rotating 4.8

Figure 4-11. Fuselage and fin for directional stability.

to the right, the air is striking the left side of the fin at an angle. Tis causes pressure on the left side of the fin, which resists the turning motion and slows down the aircraft’s yaw. In doing so, it acts somewhat like the weather vane by turning the aircraft into the relative wind. Te initial change in direction of the aircraft’s flightpath is generally slightly behind its change of heading. Terefore, after a slight yawing of the aircraft to the right, there is a brief moment when the aircraft is still moving along its original path, but its longitudinal axis is pointed slightly to the right. Te aircraft is then momentarily skidding sideways, and during that moment (since it is assumed that although the yawing motion has stopped, the excess pressure on the left side of the fin still persists) there is necessarily a tendency for the aircraft to be turned partially back to the left. Tere is a momentary restoring tendency caused by the fin. Tis restoring tendency is relatively slow in developing and ceases when the aircraft stops skidding. When it ceases, the aircraft is flying in a direction slightly different from the original direction. It will not return of its own accord to the original heading; the pilot must reestablish the initial heading. A minor improvement of directional stability may be obtained through sweepback. Sweepback is incorporated Module 08 - Basic Aerodynamics

in the design of the wing primarily to delay the onset of compressibility during high-speed flight. In lighter and slower aircraft, sweepback aids in locating the center of pressure in the correct relationship with the CG. A longitudinally stable aircraft is built with the center of pressure aft of the CG. Because of structu ral reasons, aircraft designers sometimes cannot attach the wings to the fuselage at the exact desired point. If they had to mount the wings too far forward, and at right angles to the fuselage, the center of pressure would not be far enough to the rear to result in the desired amount of longitudinal stability. By building sweepback into the wings, however, the designers can move the center of pressure toward the rear. Te amount of sweepback and the position of the wings then place the center of pressure in the correct location. he contribution of the wing to static direc tional stability is usually small. he swept wing provides a stable contribution depending on the amount of sweepback, but the contribution is relatively small when compared with other components.


FREE DIRECTIONAL OSCILLATIONS (DUTCH ROLL) Dutch roll is a coupled lateral/directional oscillation that

is usually dynamically stable but is unsafe in an aircraft because of the oscillatory nature. Te damping of the oscillatory mode may be weak or strong depending on the properties of the particular aircraft. If the aircraft has a right wing pushed down, the positive sideslip angle corrects the wing laterally before the nose is realigned with the relative wind. As the wing corrects the position, a lateral directional oscillation can occur resulting in the nose of the aircraft making a figure eight on the horizon as a result of two oscillations (roll and yaw), which, although of about the same magnitude, are out of phase with each other. In most modern aircraft, except high-speed swept wing designs, these free directional oscillations usually die out automatically in very few cycles unless the air continues to be gusty or turbulent. Tose aircraft with continuing Dutch roll tendencies are usually equipped with gyrostabilized yaw dampers. Manufacturers try to reach a midpoint between too much and too little directional stability. Because it is more desirable for the aircraft to have "spiral instability" than Dutch roll tendencies, most aircraft are designed with that characteristic.

4.9


4.10


QUESTIONS

Question: 4-1 Name the three axes of an aircraft.

Question: 4-5 An increase in wing __________________ increases the lateral stability of the aircraft in flight.

Question: 4-2 Static longitudinal stability of an aircraft depends on what three things?

Question: 4-6 Stability about the vertical axis of an aircraft is known as __________________ or __________________ stability.


4.11

ANSWERS

Answer: 4-1 longitudinal. lateral. vertical.

Answer: 4-3 dihedral.

Answer: 4-2 Location of the wing with respect to the CG (center of gravity). Location of the horizontal tail surfaces in relation to the CG. Area of the tail surfaces.

Answer: 4-4 yawing. directional.

4.12


ACRONYM INDEX (ACRONYMS USED IN THIS MANUAL)

AMC AOA CD CD CG CL CL CL-MAX CP DP EASA GM Gs ICAO ISA ISO KCAS L/D L/D MA X MAC RO UV VMO/MMO

Module 08 - Acronym Index

/ / / / / / / / / / / / / / / / / / / / / / /

Acceptable Means of Compliance Angle of Attack Drag Coefficient Coefficient of Drag Center of Gravity Lift Coefficient Center of Lift Maximum Coefficient of Lift Center of Pressure Parasite Drag European Aviation Safety Administration Guidance Material Acceleration of Gravity International Civil Aviation Organization International Standard Atmosphere International Organization for Standardization Knots Calibrated Air Speed Lift/Drag Lift to Drag Ratio Mean Aerodynamic Chord Rate of Return Ultraviolet Rays Maximum Operating Limit Speed

A.1

A.2

Module 08 - Acronym Index

GLOSSARY

Acceleration . Increase in the rate or speed of something.

Critical Angle. When the critical angle is reached, the

Aerodynamics . Te study of the properties of moving

air ceases to flow smoothly over the top surface of the airfoil and begins to burble or eddy. Tis means that air breaks away from the upper camber line of the wing.

air, and especially of the interaction between the air and solid bodies moving through it.

Densit y. Te density, or more precisely, the volumetric Airfoil . A structure with curved surfaces designed to

mass density, of a substance is its mass per unit volume.

give the most favorable ratio of lift to drag in flight, used as the basic form of the wings, fins, and horizontal stabilizer of most aircraft.

Dihedral . Te wings on either side of the aircraft join

Angle of Attack . Is the angle between the oncoming air

Directional Stability. Directional stability is associated

or relative wind and a reference line on the airplane or wing.

with the realigning of the longitudinal axis with the flight path (the angle of zero slip) after a disturbance causes the aircraft to yaw out of alignment and produce slip.

the fuselage to form a slight V or angle.

Angle of Inc ide nce . he acute angle the wing chord

makes with the longitudinal axis of the aircraft is called the angle of incidence, or the angle of wing setting. At mospheric Pressure . he pressure exerted by the

weight of the atmosphere, which at sea level has a mean value of 101,325 pascals (roughly 14.6959 pounds per square inch).

Drag Rise . Te drag divergence Mach number (not to be

confused with critical Mach number) is the Mach number at which the aerodynamic drag on an airfoil or airframe begins to increase rapidly as the Mach number continues to increase. Tis increase can cause the drag coefficient to rise to more than ten times its low speed value. Drag . Te force that exerts a braking action to hold the

Axis. Tere are three axes around which the airplane

moves. Tese axes all pass through the airplane's center of gravity, which is that point which is the center of the airplane's total weight.

aircraft back. Drag is a backward deterrent force and is caused by the disruption of the airflow by the wings, fuselage, and protruding objects. Dutch Roll . Dutch roll is a type of aircraft motion,

Boundary Layer . On an aircraft wing the boundary layer

is the part of the flow close to the wing, where viscous forces distort the surrounding non-viscous flow.

consisting of an out-of-phase combination of "tail wagging" and rocking from side to side. Dynamic Stability. the aircraft response over time when

Burble Point . When the AOA increases to the angle of

disturbed from a given AOA, slip, or bank.

maximum lift, the burble point is reached. Fineness Ratio. is defined as the ratio of the chord of the Center of Pressure. he point of intersection of the

airfoil to its maximum thickness.

resultant force line with the chord line of the airfoil is called the center of pressure (CP).

Flaps . Flaps are usually mounted on the wing trailing

which extends forward of the leading edge.

edges of a fixed-wing aircraft. Flaps are used to lower the minimum speed at which the aircraft can be safely flown, and to increase the angle of descent for landing.

Chord . Te Chord of a wing is the width of the wing

Form Drag . Which results from the aerodynamic

from the leading edge apex to the trailing edge.

resistance to motion due to the shape of the aircraft.

Chord Line. A Chord Line is a line depicting the chord

Module 08 - Glossary

G.1

GLOSSARY

Glide Angle. Te angle between the plane of the horizon

Mean Aerodynamic Chord . Is the average distance from

and the path of a glider or airplane; especially: the least angle at which a glider or airplane will glide to earth in still air.

the leading edge to the trailing edge of the wing. Due to the many wing planform designs, the mean aerodynamic chord is not necessarily half way from the fuselage to the wing tip as it is on a perfectly rectangular wing.

Glide Ratio. Te distance the airplane will, with power

off, travel forward in relation to the altitude it loses.

Parasite Drag . Parasite (parasitic) drag (DP) is defined as

all drag that is not associated with the production of lift. Gravity or Weight . he force that pulls the aircraft

toward the earth. Weight is the force of gravity acting downward upon everything that goes into the aircraft, such as the aircraft itself, crew, fuel, and cargo.

Planform . Te outline of wing or tailplane when viewed

from above. Polar Curve. A graph which contrasts the sink rate of an

Gravity. Is the pulling force that tends to draw all

aircraft with its horizontal speed.

bodies toward the center of the earth. Shock Wave . ypically associated with supersonic Humidity. A quantity representing the amount of water

vapor in the atmosphere or a gas.

aircraft, however, they also form on an aircraft traveling at less than the speed of sound.

Inter ference Drag . Which occurs where surfaces with

Skin Friction Drag . Which is related to the smoothness

different flow characteristics meet (e.g. wing and fuselage).

(or roughness) of the aircraft surfaces. Speed . Te rate at which someone or something is able

Lat era l Stability. Stability about the aircraft’s

to move or operate.

longitudinal axis, which extends from the nose of the aircraft to its tail.

Stall Fence. Wing fences, also known as boundary

Lift Coeff icient . Is a ratio between lift pressure and

dynamic pressure and is a function of the shape of the wing and angle of attack.

layer fences and potential fences are fixed aerodynamic devices attached to aircraft wings. Often seen on swept wing aircraft, wing fences are f lat plates fixed to the upper surfaces parallel to the airflow, typically wrapping around the leading edge.

Lift . Te force that pushes the aircraft upward. Lift acts

vertically and counteracts the effects of weight.

Stall . An aircraft stall results from a rapid decrease in

Load Factor . Load factors are also referred to as "limit

lift caused by the separation of airflow from the wing’s surface brought on by exceeding the critical AOA. A stall can occur at any pitch attitude or airspeed.

load factors." For reasons of safety, it is required that the aircraft be designed to withstand load factors without any structural damage.

Static Stability. he initial tendency to return to

Longitudinal Stabilit y. Is the quality that makes an

equilibrium that the aircraft displays after being disturbed from its trimmed condition.

aircraft stable on its lateral axis. It involves the pitching motion as the aircraft’s nose moves up and down in flight.

Subsonic . Te term describes an aircraft that flies below

its critical Mach number. Mach Number . Te ratio of the true airspeed of the aircraft

to the speed of sound in the same atmospheric conditions.

G.2


GLOSSARY

Supersonic . he term describes an aircraft t hat fl ies

above its critical Mach number. Sweepback. Te angle at which an aircraft's wing is set

back from a right angle to the body. Trust . Te force that moves the aircraft forward. Trust

is the forward force produced by the powerplant that overcomes the force of drag. ransition Point . At the transition point, the character

of the airf low changes and the laminar boundary layer quickly becomes a turbulent boundary layer. his tu rbu lent boundar y layer cont inues to thicken downstream. Velocity. Te speed of something in a given direction. Vortex Generators. Is an aerodynamic device, consisting

of a small vane usually attached to a lifting surface, such as an aircraft wing. Vortex. A mass of whirling fluid or air, especially a

whirlpool or whirlwind. Wave Drag . Is a force, or drag that retards the forward

movement of an airplane, in both supersonic and transonic flight, as a consequence of the formation of shock waves. Winglet . A small vertical surface at the tips of the wings.

hey help direct the turbulent airf low that all wings have at the tips. Tey make the wings more efficient. Yawing . A yaw motion is a side to side movement of the

nose of the aircraft. Te yawing motion is being caused by the deflection of the rudder.


G.3

G.4


INDEX

A

International Standard Atmosphere .............................. 1.6

Aerodynamic Forces in urns ........................................ 3.5 Aerodynamic Resultant ................................................. 2.10 Aerodynamics ................................................................ 2.2 Aerodynamic erms ....................................................... 2.7 Airfoil Contamination ................................................... 2.9 Airfoils ........................................................................... 2.8 Angle of Attack (AOA) ................................................. 2.12

B Basic Aerodynamics ....................................................... 1.2 Bernoulli’s Principle and Subsonic Flow ....................... 2.3 Boundary Layer and Friction Effects ............................. 2.4

D Density ........................................................................... Dihedral ......................................................................... Drag ............................................................................... Drag and Airspeed .........................................................

1.4 4.7 2.14 2.17

F

K Keel Effect and Weight Distribution ............................. 4.7

L Lateral Stability (Rolling) .............................................. 4.7 Leading Edge Devices ................................................... 3.17 Lift ................................................................................. 3.2 Lift and Drag Coefficients ............................................. 2.11 Lift Augmentation ......................................................... 3.16 Lift/Drag Ratio ............................................................. 2.11 Load Factors................................................................... 3.9 Load Factors in Aircraft Design .................................... 3.9 Longitudinal Stability (Pitching) .................................. 4.4

M Mach Number vs A irspeed ............................................ 3.12

N Newton’s Laws of Motion .............................................. 2.2

Fixed Airflow Devices ................................................... 3.18 Flaps ............................................................................... 3.16 Flight Envelope and Structural Limitations .................. 3.9 Form Drag...................................................................... 2.14 Free Directional Oscillations (Dutch Roll) ................... 4.9 Friction Drag ................................................................. 2.14

G Generation of Lift and Drag .......................................... 2.11 Glide Ratio..................................................................... 3.3

H High Speed Flight ......................................................... 3.10 Humidity ........................................................................ 1.4

I Induced Drag ................................................................. 2.15 Interference Drag ........................................................... 2.15 Module 08 - Index

P Parasite Drag .................................................................. Physics of the Atmosphere............................................. Planform and Vortices ................................................... Polar Curve .................................................................... Pressure ..........................................................................

2.14 1.2 2.7 3.5 1.2

R Relationship Between Lift, Weight, Trust and Drag .. 3.2

S Shape of the Airfoil ........................................................ Shock Waves .................................................................. Speed Ranges ................................................................. Stability and Control...................................................... Stability (Yawing) .......................................................... Stalls ..............................................................................

2.9 3.13 3.11 4.3 4.8 3.7 I.1

INDEX

Static Stability ................................................................ 4.3 Subsonic vs Supersonic Flow ......................................... 3.10 Sweepback ...................................................................... 3.14 Sweepback ...................................................................... 4.7

T emperature and Altitude.............................................. Te Axes of an Aircraft.................................................. Trust and Drag ............................................................. Trust, Weight, Lift and Drag ......................................

1.5 4.2 3.3 2.10

V Velocity and Acceleration............................................... 2.2 Vertical (Directional) ..................................................... 4.8

W Wave Drag ..................................................................... 2.16 Weight, Center of Gravity, Lift ..................................... 3.2

I.2

Module 08 - Index

basic

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